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THE

AIRCRAFT

NAVAIR 01 -HI AAB-1
PRELIMINARY

NATOPS FLIGHT MANUAL

NA VY MODEL

AH-1T (TOW)
AIRCRAFT

This publication is required for official use or for administrative or opera¬
tional purposes only. Distribution is limited to U.S. Government agencies.
Other requests for the document must be referred to Commanding Officer,
Naval Air Technical Services Facility, 700 Robbins Avenue, Philadelphia,
PA 19111.

ISSUED BY AUTHORITY OF THE CHIEF OF NAVAL OPERATIONS
AND UNDER THE DIRECTION OF THE COMMANDER
NAVAL AIR SYSTEMS COMMAND

1 AUGUST 1980

Change 1—15 February 1983

NAVAIR 01-H1AAB-1

Reproduction for non-military use of the information or illustrations contained in this publication is not
permitted without specific approval of the Commander, Naval Air Systems Command.

—--LIST OF EFFECTIVE PAGES--

Date of issue for original pages is:

Original.0.1 October 1980

Change.1 . . . 15 February 1983

Total Number of Pages in this Publication is 379,
consisting of the following:

Page No. Issue

Cover.0

*Title.1

*A-B.1

C.0

Letter of Promulgation

(Reverse Blank).0

*i.1

ii—vi.0

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The asterisk indicates pages changed, added, or deleted by the current change.

Change 1

NAVAIR 01 -HIAAB-1

INTERIM CHANGE SUMMARY

The following Interim Changes have been cancelled or previously incorporated in this manual:

INTERIM

CHANGE

NUMBER(S)

REMARKS/PURPOSE

1—6, 8

The following Interim Changes have been incorporated in this Change/Revision:

INTERIM

CHANGE

NUMBER

REMARKS/PURPOSE

Airspeed Limits.

Rotor Brake Limits.

Impending Transmission Failure.

Mast Bumping.

Maneuvering Flight.

Interim Changes Outstanding — To be maintained by the custodian of this manual:

INTERIM

CHANGE

NUMBER

ORIGINATOR/DATE
(or DATE/TIME GROUP)

PAGES

AFFECTED

REMARKS/PURPOSE

NAVAIR 01-H1AAB-1

SUMMARY OF APPLICABLE TECHNICAL DIRECTIVES

Information relating to the following recent technical directives has been incorporated in this manual

CHANGE

NUMBER

DESCRIPTION

DATE INC.

IN MANUAL

VISUAL IDENTIFICATION

Information relating to the following recent technical directives will be incorporated in a future change

CHANGE

NUMBER

DESCRIPTION

VISUAL IDENTIFICATION

NAVAIR 01 -HIAAB-1

DEPARTMENT OF THE NAVY
OFFICE OF THE CHIEF OF NAVAL OPERATIONS
WASHINGTON, D.C. 20350

LETTER OF PROMULGATION

1. The Naval Air Training and Operating Procedures Standardization Program
(NATOPS) is a positive approach toward improving combat readiness and achieving
a substantial reduction in the aircraft accident rate. Standardization, based on
professional knowledge and experience, provides the basis for development of an
efficient and sound operational procedure. The standardization program is not
planned to stifle individual initiative, but rather to aid the Commanding Officer in
increasing his unit’s combat potential without reducing his command prestige or
responsibility.

2. This manual standardizes ground and flight procedures but does not include
tactical doctrine. Compliance with the stipulated manual procedure is mandatory
except as authorized herein. In order to remain effective, NATOPS must be dynamic
and stimulate rather than suppress individual thinking. Since aviation is a continuing,
progressive profession, it is both desirable and necessary that new ideas and new
techniques be expeditiously evaluated and incorporated if proven to be sound. To
this end, Commanding Officers of aviation units are authorized to modify
procedures contained herein, in accordance with the waiver provisions established by
OPNAVINST 3510.9 series, for the purpose of assessing new ideas prior to initiating
recommendations for permanent changes. This manual is prepared and kept current
by the users in order to achieve maximum readiness and safety in the most efficient
and economical manner. Should conflict exist between the training and operating
procedures found in this manual and those found in other publications, this manual
will govern.

3. Checklists and other pertinent extracts from this publication necessary to normal
operations and training should be made and may be carried in Naval Aircraft for use
therein. It is forbidden to make copies of this entire publication or major portions
thereof without specific authority of the Chief of Naval Operations.

w. l. McDonald
V ice Admiral, USN
Deputy Chief of Naval Operations
(Air Warfare)

NAVAIR 01 -HI AAB-1

Contents

TABLE OF CONTENTS

SECTION I HELICOPTER.

Part 1 Helicopter and Engines.

Part 2 Systems...

Part 3 Service and Handling.

Part 4 Operating Limitations.

SECTION II INDOCTRINATION.

SECTION III NORMAL PROCEDURES.

Part 1 Flight Preparation.

Part 2 Shore-Based Procedures.

Part 3 Ship-Based Procedures.

Part 4 Special Procedures.

Part 5 Functional Checkflight Procedures.

SECTION IV FLIGHT CHARACTERISTICS.

SECTION V EMERGENCY PROCEDURES .

Part 1 Ground Emergencies.

Part 2 Takeoff Emergencies.

Part 3 Inflight Emergencies.

Part 4 Landing Emergencies.

SECTION VI ALL WEATHER OPERATION .

Part 1 Instrument Procedures.

Part 2 Extreme Weather Operation.

SECTION VII COMMUNICATIONS — NAVIGATION EQUIPMENT AND

PROCEDURES.

SECTION VIII WEAPONS SYSTEM.

SECTION IX FLIGHT CREW COORDINATION.

SECTION X NATOPS EVALUATION.

SECTION XI PERFORMANCE DATA.

Part 1 Standard Data.

Part 2 Takeoff.

Part 3 Climb.

Part 4 Range.

Part 5 Endurance...

Part 6 Emergency Operation.

Part 7 Special Charts.

ALPHABETICAL INDEX.

FOLDOUT ILLUSTRATIONS.

1-1

1- 7
1-56
1-66

2 - 1
3-1

3-1

3- 5
3-17
3-29
3-35

4- 1

5- 1

5-5

5-7

5- 8
5-30

6 - 1

6-1

6-2

7-1

8-1

9-1

10-1

11-1

11-10

11-15

11-18

11-49

11-55

11-87

Index-1

FO-O

Change 1 i

Foreword

NAVAIR 01 -HIAAB-1

FOREWORD

SCOPE

The NATOPS b light Manual is issued by the
authority of the Chief of Naval Operations and
under the direction of Commander, Naval Air
Systems Command in conjunction with the Naval
Air I raining and Operating Procedures
Standardization (NATOPS) Program. This
manual contains information on all aircraft
systems, performance data, and operating
procedures required for safe and effective
operations. However, it is not a substitute for
sound judgement. Compound emergencies,
available facilities, adverse weather or terrain, or
considerations affecting the lives and property of
others may require modification of the procedures
contained herein. Read this manual from cover to
cover. It’s your responsibility to have a complete
knowledge of its contents.

APPLICABLE PUBLICATIONS

The following applicable publications complement
this manual:

NAVAIR 01-H1AAB-1B (Pocket Checklist)
NAVAIR 01-H1 AAB-1 F (Functional Checkflight
Checklist)

HOW TO GET COPIES

Each flight crewmember is entitled to personal copies
of the NATOPS Flight Manual and appropriate
applicable publications.

Automatic Distribution

To receive future changes and revisions to this
manual or any other NAVAIR aeronautical
publication automatically, a unit must be
established on an automatic distribution list

maintained by the Naval Air Technical Services
Facility (NATSF). To become established on the list
or to change existing NAVAIR publication
i equirements, a unit must submit the appropriate
tables from NAVAIR 00-25DRT-1 (Naval Aeronautic
Publications Automatic Distribution Requirement
Tables) to NATSF, Code 321, 700 Robbins Avenue,
Philadelphia, PA 19111. Publication requirements
should be reviewed periodically and each time
requirements change, a new NAVAIR 00-25DRT-1
should be submitted. NAVAIR 00-25DRT-1 only
provides for future issues of basic, changes, or
revisions and will not generate supply action for the
issuance of publications from stock. For additional
instructions, refer to NAVAIRINST 5605.4 series
and Introduction to Navy Stocklist of Publications
and Forms NAVSUP Publication 2002 (S/N
0535-LP-004-0001).

Additional Copies

Additional copies of this manual and changes thereto
may be procured by submitting DD Form 1348 to
NAVPUBFORMCEN Philadelphia in accordance
with Introduction to Navy Stocklist of Publications
and Forms NAVSUP Publication 2002.

UPDATING THE MANUAL

To ensure that the manual contains the latest
procedures and information, NATOPS review
conferences are held in accordance with
OPNAVINST 3510.9 series.

CHANGE RECOMMENDATIONS

Recommended changes to this manual or other
NATOPS publications may be submitted by anyone
in accordance with OPNAVINST 3510.9 series.

NAVAIR 01-H1AAB-1

Foreword

Routine change recommendations are submitted
directly to the Model Manager on OPNAV Form
3500-22 shown on the next page. The address of the
Model Manager of this aircraft is:

Commanding Officer

HMA-269, MAG-29, MCAS (H) New River
Jacksonville, N.C. 28545
. (Attn: NATOPS)

Change recommendations of an URGENT nature
(safety of flight, etc.,) should be submitted directly to
the NATOPS Advisory Group Member in the chain of
command by priority message.

YOUR RESPONSIBILITY

NATOPS Flight Manuals are kept current through
an active manual change program. Any corrections,
additions, or constructive suggestions for
improvement of its content should be submitted by
routine or urgent change recommendation, as
appropriate, at once.

NATOPS FLIGHT MANUAL INTERIM
CHANGES

Flight Manual Interim Changes are changes or
corrections to the NATOPS Flight Manuals
promulgated by CNO or NAVAIRSYSCOM. Interim
Changes are issued either as printed pages, or as a
naval message. The Interim Change Summary page
is provided as a record of all interim changes. Upon
receipt of a change or revision, the custodian of the
manual should check the updated Interim Change
Summary to ascertain that all outstanding interim
changes have been either incorporated or canceled;
those not incorporated shall be recorded as
outstanding in the section provided.

CHANGE SYMBOLS

0 Revised text is indicated by a black vertical line in
either margin of the page, adjacent to the affected

text, like the one printed next to this paragraph. The
change symbol identifies the addition of either new
information, a changed procedure, the correction of
an error, or a rephrasing of the previous manual.

WARNINGS, CAUTIONS, AND NOTES

The following definitions apply to "WARNINGS”,
CAUTIONS”, and "NOTES” found through the
manual.

An operating procedure, practice, or
condition, etc., which may result in injury
or death if not carefully observed or
followed.

\ Mnn ' nuu ' Munu i

CAUTION

:• j;

An operating procedure, practice, or
condition, etc., which may result in
damage to equipment if not carefully
observed or followed.

NOTE

An operating procedure, practice, or
condition, etc., which is essential to
emphasize.

WORDING

The concept of word usage and intended meaning
which has been adhered to in preparing this Manual
is as follows:

iii

Foreword

NAVAIR 01 -HIAAB-1

NATOPS/TACTICAL CHANGE RECOMMENDATION

OPNAV FORM 3500/22 (5-69) 0107-722-2002 DATE

TO BE FILLED IN BY ORIGINATOR AND FORWARDED TO MODEL MANAGER

FROM (originator)

Unit

TO (Model Manager)

Unit

Complete Name of Manual/Check list

Revision Date

Change Date

Section/Chapter

Page

Paragraph

Recommendation (be specific)

J chick ir continued on iack

Justification

Signature

Rank

Title

Address of Unit or Command

TO BE FILLED IN BY MODEL MANAGER (Ket

urn to Originator)

F HOW

DATE

HI H HI NCI

in) Your Change Rcpnmmendation Hated

□ Your change recommendation dated
review conference planned for

is acknowledged. It will be held for action of the
_ to be held at _

□

Your change recommendation is reclassified URGKNT and forwarded tor approval to
_by my DIG _

/s/

MODI I MAS Mil K .

\IR( k \l I

NAVAIR 01 -HIAAB-1

Foreword

“Shall' has been used only when application of a
procedure is mandatory.

“Should" has been used only when application of a
procedure is recommended.

“May" and “need not" have been used only when
application of a procedure is optional.

“Will" has been used only to indicate futurity,
never to indicate any decree of requirement for

application of a procedure.

NAVAIR 01-H1AAB-1

NAVAIR 01 -HI AAB-1

SECTION I — HELICOPTER

Section I
Part 1

TABLE OF CONTENTS

PART 1 - HELICOPTER AND ENGINES

Helicopter . IT

Engines . 1-2

Speed Range. 1-2

Takeoff Gross Weight. 1-2

Helicopter Arrangement. 1-2

PART 2 - SYSTEMS

Power Plant. IT

Rotor System. 1-16

Transmission System.1-18

Rotor Brake.1-23

Fuel Supply System.1-23

Auxiliary Fuel System.1-27

Pressure Fueling.1-27

DC Power Supply System. 1-28

AC Power Supply System. 1-32

External Power Receptacle. 1-32

Hydraulic Power Supply System . 1-34

Flight Control System . 1-34

Stability and Control Augmentation

System (SCAS) . 1“37

Synchronized Elevator. 1-37

Landing Gear System. 1-37

Tail Skid . 1“37

Instruments .1"37

Emergency Equipment.1-41

Crew Compartment Doors
Canopy Jettison System .

Pilot Seat.

Copilot/Gunner Seat.

Shoulder Harness.

Ventilating System.

Exterior Lights.

Interior Lights.

Part 3 - SERVICE AND HANDLING

Fueling and Servicing.1-56

Engine Wash Procedures.1-56

Pressure Hot Fueling. 1-63

Pressure Fueling.1-64

Line Operations.1-64

PART 4 - OPERATING LIMITATIONS

Instrument Markings.1-66

Torque Limits.1-66

Starter Limitations .1-66

Rotor Brake Limitations .1-66

Airspeed Limits.1-66

Prohibited Maneuvers.1-66

Minimum Crew Requirements.1-66

Center of Gravity Limitations.1-71

Lateral CG Limitations.1-71

Acceleration G Limitations.1-71

1-44

1-48

1-52

PART 1 — HELICOPTER AND ENGINES

HELICOPTER.

The AH-IT is a tandem seat, two place (pilot and
copilot/gunner) twin engine attack helicopter
manufactured by Bell Helicopter Textron. It is an
aggressive, high speed helicopter designed and
built around the fighting mission. The mission
profiles completely cover the air to ground
environment with multiple weapon suppressive

fire. The helicopter is maneuverable, capable of
low altitude, high speed flight and self protection
in hostile air and ground battle situations. Its
primary mission is search and target acquisition,
reconnaissance by fire, multiple weapon fire
support and troop helicopter support. The
helicopter is capable of performing this mission
from prepared or unprepared areas, during day or
night flying and navigating by dead reckoning or
by use of radio aids to navigation.

1-1

Section I
Part 1

NAVAIR 01-H1AAB-1

ENGINES.

The T400-WV-402 engine is a twin power section
turboshaft engine consisting of two identical free-
turbine turboshaft power sections driving a single
output shaft through separate halves of a common
combining gearbox. The engine develops 1970
shaft horsepower at 100 percent torque and 100
percent RPM.

The engine is manufactured by Pratt & Whitney of
West Virginia.

NOTE

To enable the use of standard
terminology in this manual, the
individual power sections will ' be
referred to as engine 1 (left) and engine 2
(right).

SPEED RANGE.

The speed range of this helicopter, clean
configuration, is 0 to 190 knots based on standard
day conditions (29.92 inches of mercury, 15 degrees
Celsius at sea level).

TAKEOFF GROSS WEIGHT.

The maximum gross weight for takeoff is 14,000
pounds.

HELICOPTER ARRANGEMENT.

Refer to figure 1-1 for general arrangement.

Refer to figure 1-2 for principal dimensions.

Refer to figure 1-5 for pilot station diagram.

Refer to figure 1-4 for copilot/gunner station
diagram.

1-2

NAVAIR 01 -HI AAB-1

Section I
Part 1

1. HUB AND BLADE ASSEMBLY

2. RIGHT SYNCHRONIZED ELEVATOR

3. TAIL ROTOR HUB AND BLADE

4. 90 DEGREE GEARBOX

5. TAIL SKID

6. LEFT SYNCHRONIZED ELEVATOR

7. AFT ELECTRONICS COMPARTMENT DOOR

8. ENGINE COMPARTMENT (RH NOT SHOWN)

9. EXTERNAL POWER RECEPTACLE DOOR

10. ENGINE FIRE EXTINGUISHER ACCESS DOOR

11. TRANSMISSION COMPARTMENT (RH NOT SHOWN)

12. HYDRAULIC COMPARTMENT DOOR (RH NOT SHOWN)

13. FREE AIR TEMPERATURE GAGE

14. COPILOT/GUNNER DOOR SWITCH

15. AMMUNITION COMPARTMENT DOOR (RH NOT SHOWN)

16. TELESCOPIC SIGHT UNIT

17. RAIN REMOVAL DUCT

18. PITOT TUBE

19. PYLON ACCESS FAIRING

20. PYLON ACCESS DOOR (RH NOT SHOWN)

Figure 1-1. General Arrangement

1-3

17 FT 4 IN

Section I
Part 1

NAVAIR 01-H1AAB-1

210470-8

Figure 1-2. Principal Dimensions

1-4

14 FT 2 IN

NAVAIR 01 -HI AAB-1

Section I
Part 1

■

1. LEFT DEFROST (RH NOT SHOWN)

2. PILOT FIXED SIGHT

3. CANOPY DOOR SWITCH

4. SMOKE GRENADE CONTROL PANEL (NOT SHOWN)

Figure 1-3. Pilot Station

1-5

Section I NAVAIR 01-H1AAB-1

Part 1

Figure 1-4.

Copilot/Gunner Station

1-6

NAVAIR 01-H1AAB-1

Section I
Part 2

PART 2 — SYSTEMS

POWER PLANT.

The power plant consists of two independent
engines driving a combining gearbox. The
combining gearbox contains an overrunning
clutch and a torque meter for each engine. Each
engine has a three stage axial, single stage
centrifugal compressor driven by a single stage
turbine. Another single stage turbine
counterrotating with the first, drives into the
combining gearbox. Fuel is sprayed in the annular
combustion chamber by fourteen individually
removable fuel nozzles mounted around the gas
generator case. A high tension ignition unit (figure
1-7), and two spark igniter plugs are used to start
combustion. A hydro-pneumatic fuel control
schedules fuel flow to provide the power required to
maintain the desired output shaft speed.
Automatic load sharing is provided. In addition, a
manually operated fuel system is provided within
the fuel control system. Refer to figure 1-9.

Combining Gearbox.

The combining (reduction) gearbox located on the
aft portion of the engines (figure 1-6) has two
identical reduction gear trains which transmit
torque from each engine to a common output shaft.
Each gear train has three stages. A uni-directional
drive clutch is incorporated with the third stage
shaft allowing torque to be transmitted in one
direction only. A gearbox output section is
composed of the common output-shaft and both
third stage shafts. The lubrication for the gearbox
output section is independent of the engines and
self-contained with the combining gearbox. The
first two stages and accessory drives obtain
lubrication from their respective engines. A power
turbine governor tachometer-generator, oil cooler
blower pad and torquemeter oil pressure trans¬
mitter are fitted for each engine on separate
mounts on the combining gearbox and are driven
by their respective engines.

Engine Air Particle Separator System.

The particle separator consists of a series of
inter-connected ducts and valves that provide

each engine with air free of foreign particles
(figure 1-8). During normal operating conditions,
air approaching the engine inlet is partially
bypassed through an ejector. Air flow is controlled
by a two position door located downstream of the
engine inlet duct. With the door in the open
position, action of the air turning into the engine
inlet will remove foreign particles.

Particle Separator Switch.

The particle separator switches are three-position
switches labeled ON, AUTO, and OFF (figure 1-5).
The ON position opens the door and the PART
SEP OFF caution light will extinguish when the
door is completely open. The OFF position closes
the door to route all air into the engine and the
PART SEP OFF caution light will illuminate
when closed. The AUTO position ties the door
actuation in with the engine low rpm indication
system. In the AUTO position the door will be open
unless the gas producer rpm drops to 52.5 ± 2%.
When either engine drops to 52.5 ±2% gas producer
rpm, the RPM caution light will illuminate, the
particle separator door will close, and the PART
SEP OFF caution light will illuminate.

Actuation of a FIRE PULL handle will close the
respective door, regardless of switch position.
Power is supplied by 28 vdc essential bus and is
protected* by ENG AIR BYP VALVE circuit
breaker.

NOTE

The AUTO position shall be used at all
times when visible moisture is evident.

PARTICLE SEPARATOR CAUTION LIGHT.

The particle separator caution lights, one for each
engine duct, are located on the pilot caution panel
and labeled PART SEP OFF. Any position other
than full open will cause the light to illuminate.

Engine Fuel System.

The engine fuel system consists of separate
identical engine fuel control systems and fuel
pumps plus a common torque control unit (figure 1-
9), Fuel from the boost pump enters the engine fuel
pump housing and passes through a filter (with

1-7

Section I
Part 2

NAVAIR 01 -HIAAB-1

| OIL PRESS 1

| CHIP DET |t

SPARE ||

- >|

1 FILTER 1|

1 PART |l

1 SEP Of F J|

210900-1-1A

Figure 1-5. Engine Controls and Indicators (Sheet 1 of 2)

1-8

NAVAIR 01-H1AAB-1

Section I
Part 2

NOMENCLATURE
GAS PROD indicator
INLET TEMP indicator
ENGINE OIL indicator

TORQUE indicator
TACHOMETER indicator
Caution lights
Advisory lights

Circuit breakers

Throttle

START switch

IDLE STOP REL switch

RPM switch

ENG TRIM+/-

ENG FUEL switch

GOV switch

PART SEP switch

FUNCTION

Displays percent Ng rpm

Displays degrees Celsius inlet turbine temperature

Displays oil temperature in degrees Celsius/
oil pressure in psi.

Displays engine and transmission torque on respective needle.
Displays percent Nf/Nr rpm on respective needle.

Illuminate to show fault condition (yellow)

Illuminate to show advisory condition (green)

Protect engine electrical circuits
Manual fuel control of each engine
Energizes start cycle to respective engine
Moves idle stop to allow full throttle travel
INC/DECR — Adjust Nf rpm
Match engines

Supplies fuel to respective engine fuel control

Selects AUTO or MANUAL mode of fuel
control operation to respective engine

OFF — Particle separator door closed

AUTO — Particle separator door position open or
closed depending on engine rpm

ON — Particle separator door open.

210900-1-2A

Figure 1-5. Engine Controls and Indicators (Sheet 2 of 2)

bypass capability), then to the fuel pump. From the
fuel pump, fuel passes through a transfer valve to
either automatic or manual fuel control units as
selected.

AUTOMATIC FUEL CONTROL UNIT (AFCU).

The automatic fuel control system of each engine
consists of an automatic fuel control unit
(incorporating a gas producer turbine governor)
and a power turbine governor (common to both
engines automatic fuel control systems) (figure 1-
9). The automatic fuel control unit integrates
signals from the power turbine governor and from
its integral gas producer turbine governor, and

allows a fuel flow up to the maximum demand,
provided that the sensed limits are not exceeded.

POWER TURBINE GOVERNOR. The power
turbine governors are mounted on the combining
gearbox and are driven to a speed proportional to
that of the power turbine (figure 1-9). It supplies a
signal to the automatic fuel control unit to change
gas producer turbine speed whenever it detects a
power turbine speed change. When the engine is
operating in automatic mode, the automatic fuel
control unit may be set to maximum position
without incurring a power turbine overspeed. The
governing speed is adjusted by the RPM switch
(figure 1-5).

1-9

Section I
Part 2

NAVAIR 01-H1AAB-1

TORQUEMETER PRESSURE
OUTPUT TO TORQUE CONTROL

PNEUMATIC
ACCUMULATOR

BLOWER PAD

ACCESSORY PAD OIL
SEAL CAVITY DRAIN

TORQUEMETER
PRESSURE OUTPUT

BLOWER PAD

OIL PRESSURE SUPPLY
TO TORQUEMETER VALVE

ACCESSORY PAD OIL
SEAL CAVITY DRAIN

INSPECTION

PORT

T-5 LIMITER
(OPTIONAL)

ACCESSORY
LUBRICATION

TACHO DRIVE
ENGINE
OIL FILTER
STATIC-
CHECK VALVE

PRESSURE RETl
FROM FILTER

OIL INLET
FROM COOLER

CHIP DETECTOR

POWER TURBINE
GOVERNOR

PNEUMATIC

ACCUMULATOR

INSPECTION

PORT

POWER TURBINE
GOVERNOR
TACHO DRIVE

T-5 LIMITER
(OPTIONAL)

STATIC

CHECK VALVE
ENGINE OIL
FILTER

CLEANING
PORT PLUG
OIL INLET
FROM COOLER

OIL OUTLET
TO COOLER

GEARBOX CHIP
DETECTOR
(Not Shown)

PRESSURE
RETURN FROM
FILTER

212061-37-2A

Figure 1-6. Aft Sections of Engines Combining Gearbox and Accessories

MANUAL FUEL CONTROL.

In manual operation, the solenoid is energized
causing the transfer valve to direct fuel flow to the
manual fuel control. The pilot controls fuel flow by
throttle movement.

The manual control does not incorporate devices
to automatically limit the major engine parameters,
namely, Ng and Nf, torque and ITT. Consequent¬
ly in the manual mode of operation the operator
must monitor all engine instruments, make changes
in power settings gradually and exercise the
additional skill required to ensure that operating
limits are not exceeded.

Engine Controls.

THROTTLES.

The throttles are located on each collective stick.
They consist of two grip-type throttles which are
used for manually controlling fuel flow to the

engines. The throttle grips are rotated to the left to
increase and to the right to decrease engine power.
Friction is induced into the pilot throttle grips by
rotating the rings at the upper and lower ends of
the throttle grips (figure 1-5). Index markers are
provided to show pilot the throttles are at equal
setting. Idle stops prevent inadvertent engine
shutdowns. When the ENGINE 1 GOV and
ENGINE 2 GOV switches are in AUTO, the fuel
flow is automatic. In MANUAL, fuel flow is
controlled by the pilot.

Engine Chip Detectors.

The pilot and copilot/gunner caution panels each
have chip detector caution lights for each engine.
The lights are connected to magnetic plugs, which

1-10 Change 1

NAVAIR 01-H1AAB-1

Section I
Part 2

Ng TACHOMETER GENERATOR

FUEL PUMP AND FCU
STARTER GENERATOR

Ng TACHOMETER GENERATOR -

EXCITER BOX

-ACCESSORY GEARBOX '
OIL SIGHT GAGE

-OIL PUMP

■ ENGINE FORWARD
CHIP DETECTOR
(NOT SHOWN)

OIL PUMP

EXCITER BOX

ACCESSORY GEARBOX
OIL SIGHT GAGE

ENGINE FORWARD
CHIP DETECTOR
(NOT SHOWN)

212061-37-3C

Figure 1-7. Accessory Gearbox Sections and Components

are installed in the engines. When a magnetic plug
attracts enough metal particles to complete the
circuit, the ENGINE (1 or 2) CHIP DETR caution
light will illuminate to indicate the affected
engine. Power is supplied by the 28 vdc essential
bus and protected by the CAUTION LIGHTS
circuit breaker.

Power Plant Oil Systems.

There are three independent oil systems; one for
each engine and the third for the combining
gearbox output section (figure 1-10). The operation
of the oil system is completely automatic and self- 1
regulating. Each engine oil system is used in the
torquemeter system to supply an indication of

engine torque. Each system has its own integral
tank, oil level sight glass, tank filling aperture,
filter and drain plugs. The oil level sight glass and
filling aperture for each engine are located on the
accessory gearbox. The oil level sight glass, filling
aperture, and drain plug of the combining gearbox
section oil system are located on the combining
gearbox.

POWER PLANT OIL COOLING.

Oil cooling is accomplished by a separate oil cooler
for each engine and combining gearbox oil system
(figure 1-10). Each system has bypass valves for
bypassing the oil coolers. The engine oil coolers are
mounted with the transmission/ combining
gearbox oil coolers aft of the engine. The

1-11

Section I
Part 2

NAVAIR 01-H1AAB-1

PARTICLE SEPARATOR DOOR

INLET SCREEN

212061-37A

Figure 1-8. Engine Air Flow T400-WV402 Engine

combining gearbox oil cooler and transmission oil
cooler are combined in one unit, but are separate
coolers. Air for cooling the oil is provided by two
fans, mounted on and driven by the combining
gearbox. The fans automatically run when the
engine is running, and no control is provided.

Engine Idle Stop Release Switch.

The engine IDLE STOP REL switch (figure 1-5) is
a three position momentary-on type switch. The
pilot IDLE STOP REL switch is on the collective
switch box. The copilot/gunner IDLE STOP RLSE
switch is located on the miscellaneous control
panel. The switch energizes electrical solenoids
with retractable plungers. The solenoids are
mounted so that the plungers act as stops in the
throttle linkages. The stops prevent the pilot or
copilot/gunner from accidentally increasing or
decreasing throttles through idle. To open or close
throttles, the switch must be placed in ENG 1 or
ENG 2 position respectively. A five-second delay is
built into the switch to allow time to open or close
throttles. Power is supplied by 28 vdc essential bus
and protected by the IDLE STOP circuit breaker.

Engine RPM Switch.

The pilot RPM switch is mounted on the collective
switch box (figure 1-5). The copilot/gunner switch

is on the miscellaneous control panel. The pilot
switch is a five-position momentary-on type. The
forward (INC) position increases engine rpm.
The aft (DECR) position decreases engine rpm.
The INC/DECR positions control the governors
on both engines simultaneously. Regulated engine
rpm may be adjusted in flight, through the
operating range of 97% to 101.5 + 0.5% by
movement of the switch. The +/- positions trim
No. 2 engine torque to provide engine matching.
The copilot/gunner does not have trim capability.
Power is provided by the 28 vdc essential bus and
protected by the GOV CONT circuit breaker.

Droop Compensator.

Droop is defined as the speed change in engine rpm
(Nfi as power is increased from a no load condition.
It is an inherent characteristic designed into the
governor system. Without this characteristic,
instability would develop as engine output is
increased, resulting in Ng speed overshooting or
hunting the value necessary to satisfy the new
power condition. Design droop of the engine
governor system is as much as 4.5% to 5.5% (300 to
400 rpm) (flat pitch to full power). If Nf power were
allowed to droop other than momentarily, the
reduction in rotor speed would become critical;
therefore, a droop compensator is installed on the
governor control to raise Nf speed as power is

1-12

NAVAIR 01-H1AAB-1

Section I
Part 2

SOLENOID

AND

TRANSFER

VALVE

PNEUMATIC
MECHANICAL
ELECTRICAL

AUTOMATIC MODE FUEL FLOW
MANUAL MODE FUEL FLOW
1111 DELIVERY FUEL FLOW

* Torque limiter is not used in this installation.
** Tt5 limiter is not used in this installation.

-THROTTLE-

212060-21K

Figure 1-9. Engine Fuel Control System — Schematic

1-13

Section I NAVAIR 01 -HI AAB-1

Part 2

GEARBOX OIL TEMPERATURE
AND PRESSURE GAGE

OIL PRESS
TRANSMITTER

NO. 1
ENGINE

CHECK VALVE

ENGINE OIL COOLER

COMBINING GEARBOX
OIL COOLER

ENGINE OIL

COMBINING GEARBOX

j DRAIN

209062-14A

Figure 1-10. Engine Oil System

□

1-14

NAVAIR 01-H1AAB-1

Section I
Part 2

increased to the rpm value selected by the pilot.
The compensator is a direct mechnical linkage
between the collective control lever and the speed
selector lever on the Nf governor. Properly rigged,
the droop compensator will hold Nf rpm to plus or
minus 1% rpm from flat pitch climb out power.

A shear pin is incorporated in the droop
compensator linkage to permit collective control
movement in the event of a bind occurring in the
droop compensator linkage. When the pin is
sheared, the droop compensator is inoperative and
care must be taken to maneuver within power
adjustment capabilities of the governor.

Governor Switch.

The ENGINE 1 GOV and ENGINE 2 GOV (pilot),
ENG 1 GOV and ENG 2 GOV (copilot/gunner)
switches are two-position toggle switches located
on the pilot engine control panel (figure 1-5) and
the copilot/gunner miscellaneous control panel.
The AUTO position provides for automatic fuel
metering to the engines. The MANUAL (pilot),
MNL (copilot/gunner), position provides the pilot,
or copilot/gunner, with manual control of fuel flow
and illuminates the ENG 1 GOV MAN or ENG 2
GOV MAN ADVISORY light. Power is supplied
by the 28 vdc essential bus and protected by the
ENG NO. 1 GOV MNL amd ENG NO. 2 GOV MNL
circuit breakers.

Start Switch.

A three position START switch (push-down-to-
unlock type) is mounted on the pilot collective stick
switch box and is marked START/ENG 1 and 2.
The switch actuates the starter and ignition circuit
when placed in the ENG 1, or ENG 2 position.
However, the engine ignition circuit is not
energized unless the FUEL switch is ON.

The START switch is magnetically held in the
ENG 1 or ENG 2 position and shall be manually
returned to the center position. The START switch
is powered by the 28 vdc bus and is protected by
ENG NO 1 START and ENG NO 2 START circuit
breakers.

Fuel Control Line Heater.

The fuel control sense line heaters prevent
accumulation of ice in the engine governing

system by maintaining temperature of at least
40°F. The fuel control heaters are energized from
the 28 vdc essential bus, through circuit breakers
marked FUEL CONT HTR (figure 1 - 5 ). The
ENG 1 and ENG 2 FUEL switches control the
engine 1 and engine 2 heaters respectively.

Starter-Generator.

Two starter-generators are provided, one for each
engine. The starter turns the gas producer turbine
for starting.

Should an unsatisfactory start occur, observe
starter limitations (Section 1, Part 4). The starters
are operated independently by the START switch.
Power is supplied by the 24 vdc battery or from an
external 28 vdc power source plugged into the
external power receptacle. The starter-generators
are mounted on the engine accessory gearboxes.

Engine Instruments and Indicators.

Pilot engine instruments, caution panel, and
indicators are located on the instrument panel
(figure 1-5). The copilot/gunner instruments,
caution panel, and indicators are located on the
instrument panel (figure 1-4).

ENGINES AND TRANSMISSION
TORQUEMETER.

A triple torquemeter indicating torque on engines
1, 2, and transmission is located in the pilot and
copilot/gunner instrument panels (figures 1-4, 1-
5). The torquemeters indicate percentage of torque
imposed on the engines output shafts. Each
torquemeter displays the torque output of each
engine on the inner dial and the torque to the
transmission (combined torque of both engines) on
the outer dial. The torque of engine 1 is indicated
by pointer 1, the torque of engine 2 is indicated by
pointer 2, and the sum of these torques is indicated
by the cursor on the outer dial. The torquemeter
system receives power from the 26 vac essential
bus and is protected by the TRQ PRESS circuit
breaker.

INTER-TURBINE TEMPERATURE
INDICATORS.

Two inter-turbine temperature indicators, marked
INLET TEMP, are located on the pilot, and two on
the copilot/gunner, instrument panels (figures 1-4,
1-5). The indicators receive temperature

1-15

Section I
Part 2

NAVAIR 01 -HI AAB-1

indications from each engine. Each system
consists of twin leads, two bus-bars and
thermocouple probes connected in parallel. Power
is supplied by the 28 vdc essential bus and
protected by the ENG 1 ITT and ENG 2 ITT circuit
breakers. The high temperature warning light on
the face of the gage will illuminate any time 837°C
is exceeded. A BIAS TEST switch is located on the
pilot instrument panel (figure 1-5). Placing the
switch left or right deactivates the respective
engine bias and will cause inlet temperature to
show a higher indication.

ENGINE AND ROTOR TACHOMETER.

The triple tachometer is located on the pilot and
copilot/gunner instrument panels (figures 1-4, 1-
5). The indicators indicate rpm for both engines
power turbines and the main rotor. The two long
pointers marked 1 and 2 represent engines 1 and 2.
The rotor RPM indicator is marked R. The outer
scale is for engines percent rpm and the inner scale
is for rotor percent rpm. Power for each indicator
is provided by 28 vdc essential bus and protected
by the TRIPLE TACH IND circuit breaker. Normal
operation is indicated when all three points are in
synchronization.

GAS PRODUCER TURBINE TACHOMETER.

Two gas producer turbine tachometer indicators,
marked GAS PROD, are located on the pilot
instrument panel and two, marked GAS
PRODUCER, on the copilot/gunner instrument
panel (figures 1-4, 1-5). The indicators are powered
by tachometer generators geared to the engine gas
producer turbine shafts. The gas producer turbine
tachometer operates independently of the
electrical system. The indicator readings are in
PERCENT RPM.

OIL PRESSURE AND TEMPERATURE
INDICATORS.

Two engine oil pressure and temperature
indicators are located on the pilot instrument

panel (figure 1-5). There is a dual instrument for
each engine. The oil pressure indicators receive psi
indications from pressure transmitters on the
engines. The temperature indicators receive
Celsius indications from electrical resistance type
thermocouples located on the engines. The
pressure indicators are powered by 26 vac and
protected by the ENG OIL PRESS circuit breaker.
The temperature indicators are powered by the 28
vdc essential bus and protected by the ENG OIL
TEMP IND circuit breaker.

LOW OIL PRESSURE CAUTION LIGHT. The
pilot and copilot/gunner caution panels each have
ENG 1 and ENG 2 OIL PRESS CAUTION lights
for each engine. The caution lights are activated
by low pressure switches which make contact
when oil pressure drops below safe limits.

ROTOR SYSTEM.

Main Rotor.

The main rotor hub and blade assembly is a two-
bladed, semi-rigid seesaw type with preconing and
underslinging to optimize dynamic stability
(figure 1-11).

The main rotor hub and blade assembly consists of
a blade attached to each grip and spindle
assembly. The grip and spindle assembly is
attached to a common yoke assembly.

The grip and spindle assembly is the pitch change
element and consists of oil lubricated roller
bearings, elastomeric oil seals, tension torsion
straps, strap pins and fittings, spindle, grip drag
brace, and pitch horn.

The yoke assembly consists of a flex beam yoke,
with a trunnion and elastomeric bearings
mounted in the center section to form the flapping
axis 90 degrees to the pitch change axis.

The main rotor control system consists of a
swashplate mounted on a spherical surface for
cyclic input, a sleeve for collective input and
scissors levers mounted on the sleeve assembly
box for mixing these motions. Pitch links are
attached between each scissor lever and rotor
pitch horn for collective and cyclic control.

1-16

NAVAIR 01-H1AAB-1

Section I
Part 2

2X0010-31

1 .

DRAG BRACE

ANTIDRIVE LINK

BLADE PIN LOCK

DRIVE LINK

PITCH HORN

10.

COLLECTIVE LEVER

ROTOR HUB TRUNNION

11.

LEFT ROTOR BRAKE

MAIN ROTOR RETAINING NUT

12.

SWASHPLATE

GRIP RESERVOIR

13.

SCISSORS ASSEMBLY

PITCH CHANGE TUBE

Figure 1-11. Main Rotor System

Section I
Part 2

NAVAIR 01 -HI AAB-1

RPM Caution System.

The RPM caution system provides indication of
high rotor rpm, low rotor rpm and low gas producer
turbine rpm. Main components are ROTOR RPM
switch and RPM caution light (figure 1-14).
Electrical power is provided by the 28 vdc essential
bus and is protected by the RPM WRN lights
circuit breaker.

RPM CAUTION.

The RPM caution lights, located on the center of
pilot and copilot/gunner glareshields, illuminate
when rotor rpm increases to 103±2%, rotor rpm
decreases to 92±2% or either gas producer rpm
decreases to 52.5 ± 2%. When either gas producer
rpm decreases to 52.5 ± 2%, respective particle
separator door closes and PART SEP OFF caution
light illuminates.

ROTOR RPM SWITCH.

The ROTOR RPM switch, marked AUDIO and
OFF, is located on the pilot POWER panel (figure
1-14). OFF position prevents audio function during
engine starting when audio might be
objectionable. Switch automatically positions to
AUDIO when engine and rotor reach operating
rpm. AUDIO position provides an audio signal in
pilot and copilot/gunner headset when rotor rpm
decreases to 92±2%.

Tail Rotor.

The tail rotor hub and blade assembly is a two-
bladed, semi-rigid rotor with a skewed flapping
axis with preconing and underslinging to optimize
dynamic stability.

The tail rotor hub and blade assembly consists of a
blade attached to grip plates by bolts. The grip
plates are mounted on a common flex beam yoke
by a spherical pitch change bearing. A split
trunnion is mounted on the yoke center section by
spherical flapping bearings.

The tail rotor control system consist of a walking
beam and idlers for translating pedal input
through the fixed control system to the control
tube (FO-5). The control tube extends through the
90 degree gearbox and is attached to the

crosshead. Pitch change links connect the
crosshead and pitch horns for pitch changes
resulting from pedal movement. An active counter
balance system, activated by pitch change, offsets
the blades restoring moment, resulting in the
ability to fly boost-off if the requirement should
occur.

TRANSMISSION SYSTEM.

The transmission is mounted forward of the
engine and is coupled to the engine by a driveshaft.
It transmits engine power to the rotors and
accessories. The transmission system includes the
main rotor transmission systems, tail rotor
transmission system, accessory drive pads, speed
sensors, and associated lubrication systems.

Main Rotor Transmission System.

The main rotor transmission system consists of
the main transmission, mast assembly, and input
driveshaft (figure 1-12). The main transmission
accepts inputs from the engine and decreases
speed to the main rotor mast, tail rotor, and
transmission mounted accessories. The mast
assembly transmits power from the transmission
system to the main rotor. The driveshaft transmits
power from the engine to the main transmission
input quill.

TRANSMISSION OIL SYSTEM.

The transmission oil system is a wet sump type
consisting of a pressure pump, oil cooler,
automatic emergency oil cooler bypass system,
pressure relief valve and bypass manifold, oil
filter, jets, valves, and associated hardware (figure
1-13). These components are integral to the
transmission except for the oil cooler and filter
which are fuselage mounted. Transmission
lubrication is accomplished by a self-contained
pressure oil system. The oil pump is immersed in a
wet sump at the lower end of the transmission.

TRANSMISSION OIL COOLER.

The transmission oil cooler is a self-contained
system with independent thermostatic valves and
bypass provisions as a part of the transmission oil
cooling system. The system has an automatic
emergency oil cooler bypass valve that routes the
oil around the oil cooler or lines, if the cooler or
lines are ruptured.

1-18

NAVAIR 01 -HI AAB-1

Section I
Part 2

1. MAST NUT

2. MAST ASSEMBLY
CHIP DETECTORS (5)

4. INPUT DRIVE QUILL

5. PYLON FIFTH MOUNT

6. PYLON MAIN MOUNT

7. FIFTH MOUNT SUPPORT BEAM

8. SUPPORT CASE

9. SUMP

10. LIFT LINK

11. HYDRAULIC PUMP & TACHOMETER

12. ELECTRICAL HARNESS
HYDRAULIC PUMP QUILL
ROTOR BRAKES (2)

FILLER CAP

210040-72

Figure 1-12. Main Transmission

1-19

Section I
Part 2

NAVAIR 01-H1AAB-1

CAUTION ANNUNCIATOR

Figure 1-13. Transmission Oil System

1-20

NAVAIR 01 -HI AAB-1

Section I
Part 2

NOMENCLATURE

Tachometer indicator
ROTOR BRAKE warning light
RPM caution light
Circuit breakers
Rotor brake handle
ROTOR RPM switch

FUNCTION

Display rotor rpm in percent.

Illuminate to indicate rotor brake applied.
Illuminate to indicate high or low rpm.
Provide circuit protection.

Slow or hold rotor.

Deactivate audio rpm warning.

210900-111

Figure 1-14. Rotor System Indicator

1-21

Section I
Part 2

NAVAIR 01 -HI AAB-1

ACCESSORY DRIVE PADS.

The main rotor transmission system incorporates
drive pads for hydraulic pumps, rotor brake, and
rotor tachometer generator.

Transmission Indicators.

The transmission indicators consist of an oil
pressure indicator, oil pressure cautipn light, oil
temperature indicator, oil temperature caution
light, chip detector caution light and an oil bypass
caution light.. The combining gearbox indicators
consist of oil temperature and pressure indicators,
oil temperatue and pressure caution lights, and a
chip detector caution light (figures 1-4, 1-5).

OIL PRESSURE TEMPERATURE INDICATOR.

The transmission oil temperature and oil pressure
indicators are contained in the same case. It is
located in the pilot instrument panel and is
marked XMSN OIL T/P. The oil pressure indicator
is powered by the 26 vac ESSENTIAL BUS, and
protected by the XMSN OIL PRESS circuit
breaker. An electrical thermobulb transmits the oil
temperature to the indicator. The temperature
indicator is powered by 28 vdc essential bus, and is
protected by the XMSN OIL TEMP IND circuit
breaker.

OIL PRESSURE CAUTION LIGHT.

A XMSN OIL PRESS caution light is located on
the pilot and copilot/gunner caution panels. The
lights are connected to a transmission mounted
pressure switch. A drop in oil pressure below safe
operating limits illuminates the caution lights.

OIL TEMPERATURE CAUTION LIGHTS.

A XMSN OIL HOT caution light is located on the
pilot and copilot/gunner caution panels. The
lights are connected to a transmission mounted
thermoswitch. When the transmission oil
temperature is above safe operating limits, the
switch closes and the caution light illuminates.

TRANSMISSION CHIP DETECTOR CAUTION
LIGHTS.

The illumination of the XMSN CHIP DETR
caution light, on the pilot and copilot/gunner
caution panels indicates the transmission
mounted chip detectors have collected enough chip

or foreign material to complete the circuit. The
XMSN CHIP DET panel, located in the left
hydraulic compartment has five transmission
chip detector lights which isolate each of the chip
detectors. The XMSN CHIP DET panel lights are
UPPER MAST, PLNTY LS, PLNTY RS, SUMP
LS, and SUMP RS. One or more of the five lights
will be illuminated when the pilot and
copilot/gunner XMSN CHIP DETR caution light
is illuminated.

OIL BYPASS CAUTION LIGHT.

The XMSN OIL BYP caution light is illuminated
when the transmission oil system bypass valve is
in the bypass position. The transmission oil is then
being routed around the oil cooler. The bypass
valve closes automatically because of differential
flow between the pump and cooler outlet.

Combining Gearbox Oil Temperature and
Pressure Indicator.

The combining gearbox oil temperature and
pressure indicator is a dual indicator, registering
temperature in Celsius and pressure in psi (figure
1-36). The temperature portion receives indications
from an electrical resistance bulb and the pressure
portion receives its signal from the pressure
transmitter. The temperature portion is powered
by the 28 vdc essential bus and protected by the C
BOX OIL TEMP IND circuit breaker. The pressure
portion is powered by the 26 vac essential bus and
protected by the C BOX OIL PRESS circuit
breaker. If oil pressure in the gearbox drops below
safe limit, the C BOX OIL PRESS caution light
will illuminate. The caution light is powered by the
28 vdc essential bus and protected by the
CAUTION LIGHTS circuit breaker.

COMBINING GEARBOX CHIP DETECTOR
CAUTION LIGHT.

The illumination of the C BOX CHIP DETR
caution light on the pilot and copilot/gunner
caution panels indicates the combining gearbox
chip detector has collected enough chips or foreign
material to complete the circuit.

Tail Rotor Transmission System.

The tail rotor transmission system consists of
shaft assemblies, hanger bearing assemblies,
flexible couplings, and gearboxes. Four hanger

1-22

NAVAIR 01-H1AAB-1

Section I
Part 2

bearings are installed to support the shaft
sections. Couplings are used at the main
transmission output drive and at the forward side
of the first hanger bearing to accommodate
motion, and on the 42 degree gearbox output drive
to accommodate fin deflection. Flexible disc
couplings are installed at the remaining hanger
bearings, the 42 degree gearbox, and the 90 degree
gearbox to accommodate airframe deflections. A
fan is mounted on the 42 degree gearbox output
coupling to provide cooling for the gearbox.

42 DEGREE GEARBOX CAUTION LIGHTS.

The 42 degree gearbox caution light is located in
the pilot caution panel. The caution light will
illuminate 42° TEMP/PRESS when the oil
temperature is high or the oil pressure is low.
Illumination of the 42° CHIP DETR CAUTION
lights, located on the pilot and copilot/gunner
caution panels, indicate gearbox mounted chip
detector has collected enough chips or foreign
material to complete the circuit.

90 DEGREE GEARBOX CAUTION LIGHTS.

The 90 degree gearbox caution light is located in
the pilot caution panel. The caution light will
illuminate 90° TEMP/PRESS when the oil
temperature is high or the oil pressure is low.
Illumination of the 90° CHIP DETR CAUTION
lights, located on the pilot and copilot/gunner
caution panels, indicate gearbox mounted chip
detector has collected enough chips or foreign
material to complete the circuit.

ROTOR BRAKE.

The rotor brake is provided for stopping rotation of
the main rotor after engine shutdown and for
holding the rotor blade from turning for single
engine start. The rotor brake system consists of a
rotor brake handle, connecting cable, rotor brake
control unit, hydraulic lines, pucks, brake discs,
pressure switch, and warning light. The rotor
brake lever is located left of the pilot collective
stick (figure 1-14). The rotor brake control unit
mounted forward of the transmission is controlled
by a cable from the rotor brake handle. Brake disc
and pucks are on each side of the main
transmission. Hydraulic pressure to actuate the
rotor brake is supplied from the No. 2 hydraulic
system when the rotor is turning or by raising the
rotor brake handle through the detent gate by
pushing the detent gate button and pumping the
handle to apply brake pressure to lock the rotor.
When hydraulic pressure is applied, the pressure

switch closes the circuit and illuminates and
ROTOR BRAKE warning light on the pilot glare
shield. Power is supplied by the 28 vdc essential
bus and protected by the ROTOR BK circuit
breaker.

Operation of Rotor Brake.

To apply the brake with the rotor turning, pull
back on the rotor brake handle until the desired
stopping rate is obtained. ROTOR BRAKE light
will illuminate to indicate pressure is being
applied to the rotor brake system. When rotor has
slowed to approximately one half revolution per
minute, pull handle against the detent stop to
prevent brake pressure bleed-off.

CAUTION

Do not move rotor brake handle beyond
detent to stop a turning rotor.

NOTE

Brake pressure may be varied by
returning the handle to the full off posi¬
tion and reapplying. Maximum braking
pressure available decreases with a
decrease in rotor rpm. Small rotor brake
handle movements around intermediate
setting may result in erratic rotor brake
response.

To apply rotor brake for rotor hold during engine
start, push detent button and raise handle above
detent gate, then release detent button. Pump
handle to obtain rotor brake pressure. ROTOR
BRAKE light will illuminate to indicate that
pressure is being applied to the rotor brake system.
Continue pumping for a minimum of six strokes
after light is illuminated to assure adequate pres¬
sure is obtained.

To release rotor brake, push detent button and
place handle against the forward stop. Rotor will
start to turn when the handle is moved slightly
forward of detent gate.

FUEL SUPPLY SYSTEM.

The fuel system consists of two interconnected
rubber fuel cells. Each cell has a sump, drain valve
and a submerged fuel boost pump. In addition the
system has firewall shutoff valves, crossfeed

1-23

Section I
Part 2

NAVAIR 01-H1AAB-1

valve, low level switches, fuel feed line check
valves, boost pump pressure switches, fuel
quantity transmitters and indicator, fuel pressure
transmitter and indicator, filters, fuel cell
interconnect valve, fittings, and connecting lines
(FO-1). The crossfeed valve allows both engines to
operate from either or both fuel cells.

Engine Driven Fuel Pumps.

A pump is located on the front face of each
accessory gearbox, between the fuel control and
gearbox cover of each engine. These pumps deliver
fuel to the fuel controls. A filter is incorporated in
each engine driven fuel pump.

Fuel Switch Engine 1 and Engine 2.

Fuel switches for ENG 1 and ENG 2 are located on
the pilot ENGINE control panel (figure 1-15). The
two-position (ON-OFF) switches are lock-in-on
which must be pulled up before switch movement
to OFF can be accomplished. Movement of either
switch forward to ON energizes both fuel boost
pumps. However, each switch must be ON to open
the respective fuel valve and energize the
respective igniter plugs (the START switch on the
pilot collective stick must also be in ENG 1 or ENG
2 position to energize the respective igniter plugs).
The ON position of each FUEL switch also
energizes the fuel control heater of that engine fuel
system. The OFF position of each FUEL switch
turns off all respective fuel circuits except the fuel
boost pumps. Both switches must be in the OFF
position to turn off fuel boost pumps. Power for
all fuel circuits is supplied by the 28 vdc essential
bus and protected by the FUEL circuit breakers.

FUEL BOOST PUMPS.

One fuel boost pump is located in each fuel cell. The
pumps are electrically operated and controlled by
the ENGINE FUEL switches. Power is supplied
by 28 vdc essential bus and protected by the FWD
BOOST and AFT BOOST circuit breakers.

Fuel Interconnect Valve Switch.

The FUEL TANK INTCON switch is located on
the pilot ENGINE control panel (figure 1-15). The
OPEN position opens a valve in the connecting
line between the fuel cell. The CLOSE position
closes the valve so that no fuel can pass from one

cell to the other. Power is supplied by the 28 vdc
essential bus and protected by the FUEL VALVE
circuit breaker.

NOTE

The interconnect valve is opened when
either AUX FUEL switch is in PUMP,
regardless of FUEL INTCON switch
position.

Crossfeed Valve Switch.

The CROSSFEED valve switch is located on the
pilot ENGINE control panel (figure 1-15). When
the valve is open, either cell can supply both
engines. With the valve closed, the forward cell
supplies engine 1 only and the aft cell supplies
engine 2 only. Power is supplied by the 28 vdc
essential bus and protected by the FUEL VALVE
circuit breaker.

Fuel Quantity Indicator.

The fuel quantity indicator (figure 1-15) is located
on the pilot instrument panel. This instrument
indicates the quantity of fuel in both internal cells
in pounds, and is connected to fuel quantity
transmitters in the forward and aft cells. Power is
supplied by the 115 vac essential bus and is
protected by the FUEL QTY circuit breaker.

FUEL GAGE TEST SWITCH.

A push-button momentary-on FUEL GA TEST
switch is located on the pilot instrument panel
(figure FO-6). The switch provides a means of
testing the fuel quantity indicator and circuit for
operation. When the switch is depressed and held
in, the fuel quantity indicator pointer moves from
the actual quantity reading toward a lower
quantity reading. Upon release of the test switch,
the indicator needle will return to the actual
reading.

Fuel Pressure Indicator.

The fuel pressure indicator is located on the pilot
instrument panel (figure 1-15). The indicator
provides psi reading of fuel as delivered from the
cell mounted fuel boost pumps. The indicator is
connected to a pressure transmitter which
transmits the fuel pressure reading to the fuel
pressure indicator. With the crossfeed valve closed,
only aft boost pump pressure is indicated. Power
is supplied by the 26 vac essential bus and protected
by the FUEL PRESS circuit breaker.

1-24

LU zo

NAVAIR 01 -HI AAB-1

Section I
Part 2

FUEL*

Tty

nsr TANK CROSS

ENG 1 INTCON FEED ENG 2

ON OPEN OPEN ON

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210900-115-1

Figure 1-15. Fuel System (Sheet 1 of 2)

1-25

Section I
Part 2

NAVAIR 01-H1AAB-1

NOMENCLATURE

FUNCTION

FUEL QTY indicator
FUEL PRESS indicator
Caution lights
Circuit breaker
AUX FUEL switches

Advisory lights

Display fuel in pounds.

Display pressure in psi.

Illuminate to show fault condition.

Protect electrical circuits.

PUMP — activate respective transfer air pressure pump and signal
interconnect valve open.

OFF — deactivate respective transfer air pressure pump and signal
interconnect valve closed.

EMPTY — indicate tank is empty (yellow).

XFR — indicate fuel being transferred to forward cell (green).

ENG 1/2 switches

CROSS FEED switch

TANK INTCON switch

ENG — activate fuel system to respective engine.

OFF — deactivate fuel system to respective engine.

OPEN — allow either cell to supply either engine.

CLOSED — forward cell supplies #1 engine, aft cell
supplies #2 engine.

OPEN — allow fuel to flow between forward and aft
cell.

CLOSED — prevent fuel from flowing between
forward and aft cell.

210900-115-2

Figure 1-15. Fuel System (Sheet 2 of 2)

1-26

NAVAIR 01-H1AAB-1

Section I
Part 2

Fuel System Caution Lights.

The pilot fuel system caution lights consist of
(ENG 1) FUEL FILTER, (ENG 2) FUEL FILTER,
FWD FUEL BOOST, AFT FUEL BOOST, FWD
FUEL LOW, and AFT FUEL LOW. The gunner
fuel system caution lights consist of FWD FUEL
LOW, AFT FUEL LOW, ENG 1 FUEL FLTR and
ENG 2 FUEL FLTR.

Forward and Aft Fuel Boost Pump Caution
Lights.

The FWD FUEL BOOST and AFT FUEL BOOST
caution lights are located on the pilot caution
panel. Failure of a fuel boost pump illuminates the
caution light for that particular boost pump.

NOTE

Failure of the aft fuel boost pump, with
the fuel crossfeed valve closed, will
result in a zero fuel pressure indication
due to the fuel pressure transmitter
being located in the fuel line from the aft
cell.

Fuel Filter Caution Lights.

Caution lights for (ENG 1) FUEL FILTER and
(ENG 2) FUEL FILTER are located in the pilot
and copilot caution panel. Differential pressure
switches are mounted in the fuel lines across each
filter. When the filter becomes partially obstructed,
the pressure switch senses this and closes contacts to
energize the circuit and the (ENG 1) FUEL
FILTER caution light, or the (ENG 2) FUEL
FILTER caution light illuminates. This indicates
a partially clogged filter and impending bypass
condition. If clogging continues, the fuel bypass
valve opens to allow fuel to bypass the clogged
filter.

Forward and Aft Fuel Low Caution Lights.

Both pilot and copilot/gunner caution panels have
FWD FUEL LOW and AFT FUEL LOW caution
lights. Both fuel cells have low level switches
which illuminate the caution lights when fuel in
the cell reaches a low level. The quantity of fuel in
each cell at the time a low level light illuminates
depends on the flight attitude.

At 7 degrees nose down, (cruise attitude) with
interconnect valve open, the AFT FUEL LOW

light illuminates when 475 pounds of fuel remains,
and the FWD FUEL LOW light illuminates when
140 pounds remain. With the FUEL TANK
INTCON switch in the CLOSE position, these
values are 625 and 160 pounds respectively.

NOTE

Nose down attitude of greater than 7
degrees will result in an AFT FUEL
LOW light at higher total fuel remaining
indication.

AUXILIARY FUEL SYSTEM.

There are three types of external auxiliary fuel f
tanks which may be installed at the wing stores G
stations of the AH-1T and AH-1T (TOW) heli- I
copters. These tanks, along with associated l
accessory equipment mounted on the parent rack, I
are used in conjunction with the controls, fuel cells E
(tanks), advisory lights, and fuel and air distri- |
bution valves inside the aircraft. An automatic I
level sensing system in the aircraft forward fuel |
cell operates the air compressor (located either on I
the pylon assembly or the fuel adapter) which I
forces fuel out of the auxiliary tanks and into the
forward fuel cell. (See figure FO-1.)

Various accessories of attaching brackets and hard- I
ware, fairings, fuel and air hoses, electrical cables,
air compressors, check valves, and pressure
regulators are provided to adapt the three different
tanks to the particular aircraft and mounting
location on the aircraft. The system is designed so
that it is possible to accommodate all of the three
tank systems with only minor alterations to the
aircraft at the time of initial installation.

There are two types of 100 gallon tanks and one
type of 77 gallon tank. Both 100 gallon types
(part numbers 206K 68510-1 and 382-685001)
must be installed in pairs at the outboard positions.
The 77 gallon may be installed in either the
inboard or outboard position, in pairs, or in all
four positions.

The system using part number 206K 68510-1
(100-gallon tank), is comprised of an adapter and
pylon assembly which includes the following
components: an electrical air compressor,

an air check valve, an air pressure regulator, all of
which are installed in the pylon; fuel, air, and
electrical quick disconnects to the tank for jettison
separation, also installed in the pylon, and a tank
empty switch installed inside the tank.

Change 1 1-27

Section I
Part 2

NAVAIR 01-H1AAB-1

The systems using part number 382-68500-1 (100-
gallon tank) or part number 386-68500-1 (77-
gallon tank) is comprised of a fuel adapter and
bracket arrangement which includes the following
components: an electrical air compressor, air check
valve, and an air pressure regulator, all of which are
installed in the fuel adapter; and fuel, air, and
electrical quick disconnects to the tank for jettison
separation, also installed in the fuel adapter; and a
tank empty switch installed inside the tank.

A cockpit mounted PUMP/OFF switch is provided
for each auxiliary tank. When the switch is placed
in PUMP, fuel will transfer from the auxiliary tank
to the forward fuselage tank.

A crossfeed system is provided in the air
pressurization system so that one air pump can
pressurize both auxiliary fuel tanks if a failure of
one pump occurs. A shutoff valve is installed in the
air feed line to allow operation of only one tank.

Press-to-test lights in the cockpit indicate when
fuel is transferring and when tanks are empty.
XFR advisory indicator lights operate when fuel
is flowing through a flow indicating check valve.
EMPTY caution lights indicate when each tank is
empty.

PRESSURE FUELING.

The pressure fueling system will accept the
standard pressure fueling probe and is capable of
receiving fuel at a rate of 45 gallons per minute at
55 psi (figure 1-16). The system consists of a
receiver located in the right side of the aft fuel cell,
a dual level control valve in the forward cell, a dual
shutoff valve, and two press-to-test precheck
valves. The level control valve senses a full tank
and causes the shutoff valve to close thus stopping
the fueling operation. Proper operation of the level
control and shutoff valves are checked during
filling by pressing the precheck valves in turn
which actuates the level control valve and thus
causes the shutoff valve to close.

DC POWER SUPPLY SYSTEM.

The primary electrical power supply system is a 28-
volt direct current single-wire, negative-ground,
dual bus arrangement supplied by 30-volt, 200
ampere starter-generators, one mounted on each
engine. The system is designed so that in the event
of failure of one generator the remaining generator
will supply the electrical load. The dual-bus power
distribution system allows the nonessential loads

1 -28 Change 1

NAVAIR 01-H1AAB-1

Section I
Part 2

to be automatically deenergized in event of failure
of both generators (figure 1-17). The battery then
supplies the essential bus load. Manual selection
capability allows the nonessential bus to be
reactivated at pilot discretion. Power for turret
control and firing is supplied by the No. 2
generator when the MASTER ARM switch is in
the STBY or ARM position. Under these
conditions, power to the main bus is supplied by
the No. 1 generator. In the event of failure of the
No. 2 generator while supplying turret power, the
No. 1 generator will automatically switch to
supply turret power. The main bus is then supplied
by the battery.

DC Power Control.

The dc power is controlled by the battery switch,
generator switches, nonessential bus switch, and
dc circuit breakers.

DC POWER CONTROL PANEL.

This panel is marked POWER and is part of the
ENGINE control panel. The panel contains the
following control switches; NO. 1 GEN, NO. 2
GEN, INVERTERS, NON-ESS BUS and
BATTERY. Panel illumination is provided by a
panel light that is controlled from a rheostat
switch in the LIGHTS panel (figure 1-31).

Battery.

A 24-volt, 34 ampere-hour battery, located in the
electrical compartment, provides power for
starting when a battery start is necessary. The
battery also provides a backup source of power in
the event of generator failure. Assuming 85
percent charge, the battery can supply the
essential dc loads for a period of approximately 16
to 32 minutes depending on equipment in use.
22 vdc is minimum for battery start.

BATTERY SWITCH.

The switch is a two-position switch labeled
BATTERY (figure 1-17). Battery power is supplied
to the electrical system when the switch is in the
ON position. Placing the switch to OFF removes
battery power from the system. The
copilot/gunner is provided a two-position ELEC
PWR/EMER OFF switch on the miscellaneous
control panel to provide a means of deenergizing
the electrical system and the generator circuit.

Generator.

The starter-generators are located on the front of
each engine accessory gearbox. These are 30-volt,
200 ampere generators that deliver regulated
power when gas producer turbine rpm is
approximately 71 percent or above. Electrical
power is distributed by a dual bus arrangement.
The generator voltage regulator automatically
controls the generator field current to maintain the
proper generator output voltage of 27 to 28.5 vdc.
The generator reverse current relay automatically
opens the circuit from the generator to main dc bus
when battery voltage is greater than generator
voltage, preventing discharge of the battery
through the generator. The generator is protected
by the generator field circuit breaker.

GENERATOR SWITCHES NO. 1 AND NO. 2.

These switches are located on the pilot POWER
control panel (figure 1-17). The switches are
labeled NO. 1 GEN and NO. 2 GEN. Both switches
are normally in the ON position. The RESET
position is spring loaded to return to the OFF
position. To reset the generator, the switch must be
held in the RESET position momentarily and then
moved to the ON position. The reset circuit is
protected by the GEN BUS RESET circuit
breaker.

GENERATOR CAUTION LIGHTS.

The No. 1 DC GEN and NO. 2 DC GEN caution
lights are located on the pilot and copilot/gunner
caution panels. The lights are controlled by the
reverse current relays for each generator. When a
relay is open, the light illuminates and the
generator switch should be held in the RESET
position momentarily and then moved to ON in
an attempt to bring the generator back on the
line. When the generator starts operating, the light
will extinguish.

Nonessential Bus Switch.

The NON-ESS BUS switch is located on the pilot
POWER control panel. When the switch is in
NORMAL, power is supplied to the nonessential
bus as long as either generator is operating. In the
event of a dual generator failure, the nonessential
bus can be reclaimed by placing the switch to
MANUAL. In all normal flight operations the
switch shall be in the NORMAL position.

1-29

Section I
Part 2

NAVAIR 01-H1AAB-1

Location-Copilot/ Gunner misc. control panel

FIRE ENG 1 f—GEN NO 1-
EXT

O000000

GOV FIRE BUS

START MNL DET MAIN ITT RESET FIELD

FIRE ENG 2 f- QEN NO 2 “1
EXT

0000000

GOV FIRE BUS

START MNL DET RSV ITT RESET FIELD

IDLE GOV FORCE SAS ENG HYDR PITOT

AIR

0000000

BYP

STOP CONT TRIM PWR VALVE CONT HTR

0000000

FWD AFT CONT ANTI

BOOST BOOST VALVE HTR CKPT COLLISION NAV

00000

SRCH SRCH ROTOR RPM

PWR CONT CAUTION BK WRN

-OIL TEMP IND—|

-LIGHTS— ■ \ 0VSP

CSL

0000000

C PLT PLT PLT &

ENGXMSN BOX INSTR INSTR GUNNER GOV

0000000

TURN ALTM DC HYDR

&SUP VIB VM PRESS STBY MAIN PWR

-VOICE SECURITY-

000000

XCVR FM FM UHF UHF XCVR

AUX IFF IFF TRIPLE

TACH

UHF TACAN

0000 00

FUEL XPONDER TEST IND

ICS ICS ADF RADAR RADAR VENT LT

0000000

GUNNER

PLT GUNNER RCVR BCN ALTM BLO INSTR

210900-117-1

Figure 1-17. DC Power Supply (Sheet 1 of 2)

1-30

NAVAIR 01-H1AAB-1

Section I
Part 2

ip— f i ’.

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NOMENCLATURE

GEN switch

NON-ESS BUS

NO. 1 GEN Vf.R'iRS NO. 2 GEN
ON MA *N ON

RESET r-Tt-y RESET

QjlOFF

NON-ESS ROTOR ROrtCf.

•iVO BUS TRIM BATTERY

ws • or-R normal AursiO on on

# #e

FUNCTION

ON — activates respective generator

OFF — deactivates respective generator

RESET — flashes respective generator
field

NORMAL- removes nonessential items
from electrical circuit if generators
fail

MANUAL — allows battery operation of all
electrical circuits if generators fail

Caution lights
Circuit breakers

Illuminates to show fault condition
Protects respective circuit

NOTE

Armament DC circuit breakers
are located on the AC circuit breaker

panel. 210900 - 117-2

Figure 1-17. DC Power Supply (Sheet 2 of 2)

1-31

NAVAIR 01-H1AAB-1

DC Circuit Breaker Panel.

The dc circuit breaker panel is located on the pilot
right console (figure 1-17). Each individual circuit
breaker is labeled for the particular circuit
protected. In the event a circuit is overloaded, that
circuit breaker will pop out. The circuit is activated
by pushing the circuit breaker in.

Electrical System Indicators.

The electrical system indicators consist of a dual
ammeter and a combination ac and dc voltmeter.

DUAL AMMETER.

The dual ammeter is two ammeters in one case.
The AMPS instrument is installed in the pilot
instrument panel. The ammeters are marked 1 and
2 and indicate generator amperage output.

AC AND DC VOLTMETER. The ac and dc
voltmeters are combined in one case, marked
OLTS, and installed in the pilot instrument
panel. The left section of the voltmeter is marked
AC and represents the voltage on the 115 vac
essential bus. The right section is marked DC and
represents the voltage on the 28 vdc main bus.

AC POWER SUPPLY SYSTEM.

AC power is supplied for instrument and avionics
equipment by the 115 volt, 1,000 volt-ampere,
static main inverter. Power is supplied from the 28
vdc essential bus, but is controlled by the 28 vdc
nonessential bus. A 750 volt-ampere static inverter
is provided as a standby. The standby inverter
also receives power from the 28 vdc essential bus.
The ac power is also distributed by a dual bus
system, such that the 26 vac and 115 vac
nonessential buses are de-energized in the event of
failure of the main inverter. The standby inverter
can be manually actuated to provide power to the
ac essential buses.

AC Power Control.

The ac power is controlled by the INVERTERS
switch and the ac circuit breakers on the
ac/armament circuit breaker panel.

Inverters Switch.

The three-position (M AIN/OFF/STB Y)
INVERTERS switch is located on the pilot

POWER control panel (figure 1-18). In the MAIN
position, the main inverter is on if electrical power
is being supplied by one or both generators. The
main inverter will also be on in this position with
APU or battery power if the NON-ESS BUS switch
is in MANUAL. With the NON-ESS BUS switch in
NORMAL with both generators off and battery or
APU power applied, the standby inverter is
powering the system regardless of the
INVERTERS switch position. In the OFF
position, both inverters are off. In the STBY
position, the standby inverter is on. Under normal
conditions the switch is in MAIN. When the main
inverter is not operating, the 26 vac nonessential
bus and the 115 vac nonessential bus receive no
power.

Inverter Caution Lights.

The caution lights AC MAIN and AC STBY are
located on the pilot and copilot/gunner caution
panels. The appropriate caution light is
illuminated when ac power from either inverter to
the ac essential bus is lost.

AC/Armament Circuit Breaker Panel.

The ac/armament circuit breaker panel is located
on the pilot left-hand console (figure 1-18). Circuit
breakers of the push-pull type are pushed down to
energize and pulled up to de-energize the related
circuits. Circuit breaker switches are placed
outboard to energize and inboard to de-energize
the related circuit. In the event of a current
overload the related circuit breaker or circuit
breaker switch will be forced up or inboard and
must be reset.

EXTERNAL POWER RECEPTACLE.

The external power receptacle is located on the left
side of the fuselage. When a 28 vdc auxiliary power
unit plug is inserted in the receptacle, the external
power relay in the electrical system is energized
and 28 vdc electrical power is supplied to the
primary bus. When the external door is opened, the
EXT PWR DOOR OPEN advisory light will
illuminate on the pilot caution panel. A voltage
sensor is provided in the electrical compartment
which prevents the external power from being
supplied to the helicopter bus if the APU is not set «
within the limits of 26-29 vdc. The sensor will |
automatically disconnect the helicopter bus from “
APU if voltage moves out of limits or excessive
transients are present.

1-32 Change 1

NAVAIR 01-H1AAB-1

Section I
Part 2

ENG 28V AC FUEL FUEL TRQ

CMPS ADF

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VIB XMFR FAIL QTY PRESS PRESS XMSN ENG C BOX IND RCVR
METER

f-FORM LT-, UHF TACAN ALT RADAR GYRO ADF SCAS ATTD /

OOO0O00O0OO

PWR ROTOR FUS DF SYS ENCDR ALT CMPS IND PWR SYS
Ar r 0

IFF V€W r A?<V' REF UHF TACAN TMS

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UNr, PIT

NOMENCLATURE
Caution lights

Circuit breaker
INVERTERS switch

FUNCTION

Illuminate to show fault
condition

Protect individual circuits
MAIN — activates main inverter
OFF — deactivates inverters
STBY — activates standby inverter

210900-124

Figure 1-18.

AC Power Supply

1-33

Section I
Part 2

NAVAIR 01-H1AAB-1

HYDRAULIC POWER SUPPLY SYSTEM.

Two hydraulic systems are provided for powered
control of the cyclic, collective and directional
control system. Each system employs identical
reservoirs, modules and manifold assemblies.
Both systems are used to operate the cyclic and
collective dual power actuators in the flight control
system, there is no physical connection between
the systems. Each system uses separate passages
and piston chambers inside the dual actuators. If
one system is disabled, the other system will
supply the necessary hydraulic power for fully
normal helicopter control. Hydraulic fluid is
supplied from a pressurized reservoir to the pump.
The pumps deliver 3000 psi output pressure at
normal operating rpm. Hydraulic System No. 1
supplies system power for the cyclic, collective and
directional control actuators and the yaw SCAS
actuator. Hydraulic system No. 2 supplies system
power for the cyclic and collective control
actuators and for the pitch and roll SCAS
actuators, pylon actuators, and the rotor brake.
Hydraulic system indicators are hydraulic system
fluid level sight gages, clogged filter indicators,
pressure indicator and related system caution
light. See figure FO-2.

Hydraulic System Switch.

The hydraulic system switch is located on the pilot
POWER panel (figure 1-19). The switch is a three-
position lock-in-on switch labeled HYD SYS 1
OFF, ON and SYS 2 OFF. When the switch is in
SYS 1 OFF position, system 2 is the only system
that is supplied hydraulic pressure. When the
switch is in SYS 2 OFF position, system 1 is the
only system that is supplied hydraulic pressure.
When the switch is in the ON position, both
systems are supplied hydraulic pressure.

Fluid Level Sight Gage.

The reservoir for each system has a sight gage.
The sight gage gives a direct indication of the
hydraulic fluid level in that reservoir.

Hydraulic Filter and Indicator.

The filter assembly incorporates a red indicator
that raises when the differential pressure across the
filter element exceeds 70 psi. Once extended, the
indicator will remain so until manually depressed.
When the indicator is in the retracted (reset)
position, it is hidden from view.

Hydraulic System 1 and Hydraulic System 2
Caution Lights.

The caution panels contain segments labeled NO.
1 HYD PRESS and NO. 2 HYD PRESS.
Illumination of either segment indicates that a low
pressure condition exists in the respective system.
Segments labeled NO. 1 HYD TEMP and NO. 2
HYD TEMP are provided to indicate high
hydraulic oil temperature.

Hydraulic Gage.

The hydraulic pressure gage, marked HYD PSI, is
on the pilot instrument panel. It indicates psi
pressure of No. 1 and No. 2 hydraulic system.
Power is supplied by the 28 vdc essential bus.

FLIGHT CONTROL SYSTEM.

The flight control system is a positive mechanical
type, actuated by cyclic, collective, and tail rotor
controls. Complete controls are provided for both
pilot and copilot/gunner. The copilot/gunner
controls are slaved to the pilot controls (figures 1-
20/foldout 5). The system includes a cyclic system,
a collective control system, a tail rotor system, a
force trim system, a stability and control
augmentation system (SCAS).

Cyclic Control System.

The system is operated by the cyclic stick (figure 1-
20) movement. Moving the stick in any direction
will produce a corresponding movement of the
helicopter which is the result of a change in the
plane of rotation of the main rotor. The stick fore
and aft movement also changes the synchronized
elevator attitude to assist controllability and
lengthens eg range.

Collective Control System.

The system is operated by the collective stick
(figure 1-20). Moving the stick up or down changes
the angle of attack and lift developed by the main
rotor resulting in the ascent.or descent of the
helicopter.

Tail Rotor Control System.

The system is operated by the pedals (figure 1-20).
Pushing a pedal changes the pitch of the tail rotor

1 -34 Change 1

NAVAIR 01 -HI AAB-1

Section I
Part 2

NOMENCLATURE
HYD PSI
Caution panel
Circuit breaker
HYD switch

FUNCTION

Indicate psi of No. 1 and No. 2 system.

Illuminate to show fault condition.

Protect electrical circuit.

Controls system No. 1 or No. 2.

210900-123

Figure 1-19. Hydraulic System

1-35

Section I
Part 2

NAVAIR 01-H1AAB-1

NOMENCLATURE

FUNCTION

Cyclic
Collective
Directional pedals

Position tip path plane.
Change pitch in main rotor.
Change pitch in tail rotor.

210900-122

Figure 1-20. Flight Controls

1-36

NAVAIR 01-H1AAB-1

Section I
Part 2

resulting in directional control and may be used to
pivot the helicopter on its own vertical axis. A
pedal adjuster is provided to adjust the pedal
distance for individual comfort. Heel rests are
provided for the copilot/gunner to prevent
inadvertent pedal operation.

Force Trim System.

The system incorporates magnetic brakes and
force gradient springs in the cyclic and directional
control systems to provide artifical feel in the
systems. Depressing the cyclic stick force trim
switch will cause the magnetic brake and force
gradient to be repositioned to correspond to the
positions of the cyclic stick and pedals thus
providing trim. FORCE TRIM (Pilot), F TRIM
(copilot/gunner) switches are provided. The pilots
switch is located on the pilot POWER panel arid
the copilot/gunner switch is located on the
gunners miscellaneous control panel. The ON
position actuates the system. The OFF position of
either switch de-activates the system. Power is
supplied by the 28 vdc essential bus and protected
by the FORCE TRIM circuit breaker.

STABILITY AND CONTROL
AUGMENTATION SYSTEM (SCAS).

Description.

The SCAS is a three-axis, limited authority
stability and control augmentation system
(foldout 5). The system utilizes rate gyros and
control motion sensors to provide a well dampened
helicopter that is also responsive to pilot inputs.
The inputs are accomplished by electro-hydraulic
servo actuators installed in series with the control
system. The inputs are not felt at the pilot stick. A
hydraulic interlock automatically disengages the
appropriate channels in the event of a hydraulic
system failure.

Control Panel.

The SCAS control panel (figure 1-21) contains a
POWER switch for applying 28 vdc (essential bus)
and 115 vac operating voltages to the system. The
circuits are protected by the SAS PWR dc and
SCAS PWR ac circuit breakers. The panel also
contains three magnetically held channel engage
switches which energize electric solenoid valves
controlling hydraulic pressure to the system. The

panel has three NO-GO lights; one each associated
with PITCH, ROLL, and YAW channel engage
switches. These lights are illuminated during the
warmup to indicate the presence of current in each
associated actuator channel. Should an
engagement be attempted during this warmup
period, the actuator will make an abrupt input to
the flight controls at the moment of engagement.
When engagement is made, the NO-GO lights are
locked out of the circuit and do not operate as
malfunction indicators. Disengaging a channel,
however, restores the associated light to operation.
The NO-GO lights have a built-in press-to-test
feature for ensuring that the indicator is
operational, but this feature works only prior to
channel engagement.

SCAS (SAS) Release Switch.

The cyclic grip mounted switch (figure 1-21) is used
to disengage the pitch, roll, and yaw channels
simultaneously. The channels are re-engaged by
the PITCH, ROLL and YAW switches on the
SCAS control panel.

SYNCHRONIZED ELEVATOR.

The synchronized elevator is located near the aft
end of the tailboom and is connected by control
tubes and mechanical linkage to the fore and aft
cyclic control system (figure 1-1). Fore and aft
movement of the cyclic stick produces a change in
the synchronized elevator attitude, thus
increasing controllability and eg range.

LANDING GEAR SYSTEM.

The landing gear system is a skid type, consisting
of two lateral mounted arched crosstubes attached
to two formed longitudinal skid tubes.

TAIL SKID.

A tail skid is attached to the lower aft section of the
tailboom assembly (figure 1-1). The tail skid
reduces damage to the tailboom and tail rotor plus
acts as an indicator to the pilot in case of a tail low
landing.

INSTRUMENTS.

The flight instruments, navigation instruments
and miscellaneous instruments and indicators are
described in the following paragraphs. The

1-37

Section I NAVAIR 01 -HI AAB-1

Part 2

NOMENCLATURE

FUNCTION

NO GO indicator

POWER switch

PITCH. ROLL. YAW switches

Circuit breakers

SCAS (SAS) RELEASE button

Illuminated — Red — channel not ready for
engagement.

OFF — deactivate system.

POWER — activate system.

Up — activate respective channel.

OFF — deactivate respective channel.

Protect electrical circuit.

Deactivate pitch, roll, and yaw channels.

210900-121

Figure 1-21. Stability and Control Augmentation System (SCAS)

1-38

NAVAIR 01-H1AAB-1

Section I
Part 2

description of engine instruments, transmission
instruments and rotor instruments will be found
with the respective descriptions of the engine,
transmission and rotor. Refer to Section 7 for flight
instruments.

Airspeed Indicator.

The airspeed indicators are calibrated in knots
and provide an indicated forward airspeed. The
instrument does this by measuring the difference
between impact air pressure from the pitot tube
and static air pressure from the static vents. The
pitot tube is mounted on the pylon fairing (figure 1-
1) and the static vents are located in the side cabin
skins near the bottom edge of the canopy and just
aft of the copilot/gunner station. A pitot heater is
provided for removal of ice or snow from the pitot
tube.

Vertical Velocity Indicator.

The vertical velocity indicator indicates rate of
change of altitude in feet per minute. The
instrument is actuated by the rate of atmospheric
pressure change and is vented to the static air
system.

Altimeter.

The altimeter furnishes direct readings of height
above sea level. It is vented to the static air
pressure system and determines altitude from the
atmospheric pressure. The pilot altimeter provides
an electrical signal to the transponder for altitude
reporting.

Pilot Attitude Indicator.

The pilot attitude indicator (figure 1-22) is located
in the pilot instrument panel. The attitude
indicator provides the pilot with a visual
indication of the pitch and roll flight attitude of the
helicopter in relation to the earth’s horizontal
plane. Pitch attitude is indicated by motion of the
sphere in relation to the miniature airplane. Roll
attitude is indicated by motion of the roll pointer
with respect to the fixed roll scale located at the top
of the display. The indicator sphere can be
adjusted to zero indication by the pitch and roll
trim knobs located on the face of the instrument.
The turn and slip portion of the pilot attitude
indicator consists of a rate of turn pointer and an
inclinometer (ball) which operate independently of

each other. The electrically actuated rate of turn Q
pointer is controlled by the dc powered rate gyro. It 0
indicates in which direction and at what rate the
helicopter is turning. The inclinometer indicates
when the helicopter is balanced in flight. If the
helicopter is yawing or slipping the ball will be off
center. Just above the ball, at the six o’clock
position is a scale and a pointer which will deviate
toward the FM station when the FM control panel
mode selector switch is in the HOME position and
the FM radio is tuned to an FM station and
receiving a usable signal. When the pointer is
centered in this situation, the helicopter is on a
relative heading to or from the FM station.

This pointer will indicate TACAN course
deviation when the FM mode selector switch is in
any position other than HOME and there is a
usable TACAN signal being received by the
TACAN receiver. In this situation, the pointer will
indicate five degrees deviation from the selected
TACAN radial for each dot of pointer deflection
from center. When the pointer is centered in this
situation, the helicopter is on the selected TACAN m
radial. The flag for this pointer is at the three- Q
thirty o’clock position and will appear when the
instrument is not receiving a usable signal in its
selected mode. At the nine o’clock position on the
periphery of the instrument is a scale and a pointer
which indicates FM homing signal strength. As
the helicopter approaches the station, the pointer
will move toward the center of the scale. When the
system is not in the FM homing mode the pointer
will rest at center scale but will be accompanied by
an adjacent warning flag.

The horizontal deviation and vertical deviation
pointers are connected to FM homing. The
horizontal deviation pointer is also connected to
the homing portion of the TACAN set. The
horizontal deviation pointer is centered when the
helicopter is on course, if FM homing or TACAN is
operating. The pointer deviates to the right or left
as the helicopter moves off course. As the
helicopter approaches the station and the signal
becomes stronger the vertical deviation pointer
will move upward when the FM radio is turned to
an FM station and is in the homing mode. The
TACAN set is not connected to the vertical
deviation pointer.

Three warning flags are incorporated in the
attitude indicator. The flag centered on the left side
of the indicator is the vertical deviation warning
flag, and the flag centered on the right side of the

1-39

Section I
Part 2

NAVAIR 01-H1AAB-1

210077-32

1 PITCH TRIM

2 ROLL TRIM
3. SPHERE

4 HORIZONTAL DEVIATION FLAG
(FM AND TACAN)

5. HORIZONTAL DEVIATION POINTER
(FM AND TACAN)

6 INCLINOMETER
7. RATE-OF-TURN POINTER

8 POWER OFF FLAG

9 MINIATURE AIRCRAFT

10 VERTICAL DEVIATION FLAG (FM)

11 VERTICAL DEVIATION POINTER (FM)

Figure 1-22. Pilot Attitude Indicator

1-40

NAVAIR 01 -HI AAB-1

Section I
Part 2

indicator is the horizontal deviation warning flag.
These flags disappear from view when the
respective pointers they represent are receiving a
reliable signal. The other flag labeled OFF
appears in view when electrical power to the
instrument is off. Power is supplied by the 115 vac
essential bus and protected by the ATTD SYS
circuit breaker. Power for the integral lighting is
received from the 5 vdc lighting power supply.

Copilot/Gunner Attitude Indicator.

The attitude indicator is located on the
copilot/gunner instrument panel. This instrument
is a repeater type instrument that repeats the
information presented on the pilot instrument.
The FM homing and TACAN functions are not
connected and not functional on the
copilot/gunner attitude indicator. No turn needle
or roll trim knob is provided. Power is supplied by
the 115 vdc essential bus and protected by the
ATTD SYS circuit breaker. The integral lighting
receives power from the 5 vdc lighting power
supply.

Stand-By Compass.

A standard magnetic type compass is mounted on
the left windshield support.

Free Air Temperature Indicator.

The free air temperature indicator is located on the
left side of the pilot compartment. The indicator
provides a direct reading of the outside air
temperature.

EMERGENCY EQUIPMENT.

Pilot Master Caution System.

The pilot master caution system consists of a
segmented word CAUTION ADVISORY panel
and a remote MASTER CAUTION light.

MASTER CAUTION LIGHT.

The pilot MASTER CAUTION light is located at
the top of the glare shield. When the aviation
yellow light illuminates, the pilot is alerted to
check the caution panel for the malfunction.
Placing the caution panel switch to RESET will
extinguish and reset the MASTER CAUTION
light for subsequent caution indications.

PILOT CAUTION PANEL.

The caution panel is located in the right section of
the instrument panel. When illuminated, the
worded segment in the panel will be aviation
yellow or green (figure 1-23). When not
illuminated, the lettering will not be legible.
Illumination of any of the worded segments in the
caution panel alerts the pilot to malfunctions. The
caution panel is equipped with a RESET-TEST
switch, a BRIGHT-DIM switch and two edge
lights for illuminating the switches. Power is
supplied by the 28 vdc essential bus and protected
by the CAUTION LIGHTS circuit breaker.

BRIGHT-DIM SWITCH. The BRIGHT-DIM
switch on the pilot caution panel permits the pilot
to manually select a bright or dim condition for all
caution panel lights, MASTER CAUTION light,
ROTOR BRAKE light, RPM light, pilot armament
control, rocket control, fire warning lights in FIRE
PULL handles and RADAR ALTITUDE LOW Q
light. The intensity of the FIRE warning lights g
located on the copilot/gunner glare shield is also I
controlled by this switch. After each initial
application of power, the lamps will come on
bright. Momentarily placing the switch in the
BRIGHT position selects the bright condition
and DIM position selects dim condition.

NOTE

The dim function will operate only when

the pilot instrument lights are on.

RESET-TEST SWITCH. The pilot caution panel
has a RESET-TEST switch. The TEST position
illuminates the entire caution light system.
Testing the system will not change any
malfunction indication existing prior to testing.
The RESET position extinguishes the MASTER ■
CAUTION lights in both cockpits in preparation |
for subsequent malfunctions. ■

Copilot/Gunner Caution Panel.

The caution panel is located on the copilot instru- ■
ment panel. This is a repeater type panel and does Q
not contain as many lights as the pilot’s panel. |
The copilot has a MASTER CAUTION light I
located on the instrument glare shield. The pilot d
controls the RESET function of the copilot ■
MASTER CAUTION light. The BRIGHT-DIM I
switch and TEST switch are located on the instru- g
ment panel.

BRIGHT-DIM SWITCH. The BRIGHT-DIM switch U
enables the copilot to select a bright or dim I
condition for all caution panel lights, MASTER I
CAUTION light, RPM light and copilot armament fi
panel. ■

1-41

Section I
Part 2

NAVAIR 01 -HI AAB-1

| OIL PRESS |

f~| CAUTION- C

[ OIL PRESS |

ADVISORY

| CHIP detr|

CHIP DETR |

f\j RESET

BRIG HT |\g

SPARE

1 FUEL I

FILTER 1

UV^/yy 2

FUEL 1

FILTER 1

J TEST

DIM .

| PART 1

1 SEP OFF 1

PART 1

SEP OFF 1

| NO. 1 1

1 DC GEN

r 906 I

| TEMP/PRESS 1

f 4 " 20 ' 1

| TEMP/PRESS |

1 NO. 2

DC GEN |

| C BOX I

1 CHIP DETR 1

| XMSN 1

1 CHIP DETR 1

1 9fio |

1 CHIP DETR 1

1 1

| CHIP DETR |

kMSN

OIL HOT

f XMSN |

L_ OIL PRESS 1

( C BOX 1

1 OIL PRESS 1

| C BO X 1

OIL HOT |

| NO. 1 1

| HYD PRESS |

f AC MAIN |

| AC STBY 1

1 NO. 2 |

l HYD PRESS 1

NO. 1 1

1 HYDTEMP 1

| BATTERY I

TEMP

| XMsN I

OIL BYP 1

| NO. 2 1

1 HYD TEMP 1

| FWD

1 FUEL LOW

AFT

FUEL LOW

1 FWD

I FUEL BOOST |

1 AFT 1

| FUEL BOOST |

| ENG 1 1

1 GOV MAN

| AMMO 1

I DOOR OPEN 1

SPARE

ENG 2 II

1 GOV MAN US

ALT |

ENCODER |

EXT PWR

1 DOOR OPEN 1

.. IFF J

SPARE |fl

NO. 1

DC GEN

NO. 2

DC GEN

SPARE

ENG 1

CHIP DETR

ENG 2

CHIP DETR

XMSN

OIL PRESS

XMSN

OIL HOT

C BOX

OIL PRESS

C BOX

OIL HOT

ENG 1

GOV MAN

ENG 2

GOV MAN

AC MAIN

AC STBY

FWD

FUEL LOW

AFT

FUEL LOW

ENG 1

FUEL FLTR

ENG 2

FUEL FLTR

NO. 1

HYD PRESS

NO. 2

HYD PRESS

NO. 1

HYD TEMP

NO. 2

HYD TEMP

ENG 1

OIL PRESS

ENG 2

OIL PRESS

CHIP DETR

C BOX
CHIP DETR

XMSN

CHIP DETR

COPILOT/GUNNER

PILOT

*Only on pilot panel
**Segments aviation green

PANEL WORDING

OIL PRESS

CHIP DETR

FUEL FILTER

*PART SEP OFF

*NO. 1-2 DC GEN

*90 DEGREE TEMP/PRESS

*42 DEGREE TEMP/PRESS

C BOX CHIP DETR

XMSN CHIP DETR

90 DEGREE CHIP DETR

42 DEGREE CHIP DETR

XMSN OIL HOT

XMSN OIL PRESS

FAULT CONDITIONS

Respective engine oil pressure below operating minimum.
Metal particles in respective engine.

Fuel filter partially obstructed.

Particle separator door not full open (respective engine).
Respective DC generator failed or off.

Oil over temperature or pressure.

Metal particles in combining gearbox.

Metal particles in transmission.

Metal particles in 90 degree gearbox.

Metal particles in 42 degree gearbox.

Oil overtemperature.

Transmission oil pressure below operating minimum.

210075 - 207-1

Figure 1-23. Caution Advisory Panels (Sheet 1 of 2)

1-42

NAVAIR 01-H1AAB-1

Section I
Part 2

PANEL WORDING

C BOX OIL PRESS

C BOX OIL HOT
NO. 1-2 HYD PRESS
NO. 1-2 HYD TEMP
•BATTERY TEMP
*XMSN OIL BYP
FWD-AFT FUEL LOW
•FWD-AFT FUEL BOOST
**ENG 1-2 GOV MAN

••AMMO DOOR OPEN
••ALT ENCODER

FAULT CONDITIONS

Combining gearbox oil pressure below operating
minimum.

Oil overtemperature.

Respective hydraulic pressure below operating minimum.

Respective hydraulic fluid overtemperature.

Battery overheating.

Oil bypassing oil cooler.

Respective fuel cell quantity low.

Respective fuel boost pump pressure low.

Respective engine governor operating in manual
mode.

Ammunition compartment door open.

Electrical power lost to altimeter encoder (This light is non-
functional if the AAU-32A altimeter is installed).

“EXT PWR DOOR OPEN
“IFF

External power door open.
Kit 1A zeroized.

AC MAIN Loss of main inverter power.

AC STBY Loss of standby inverter power.

210075 - 207-2

Figure 1-23. Caution Advisory Panels (Sheet 2 of 2)

NOTE

The DIM function will operate only
when the copilot instrument lights are
on. The copilot TURRET STOW light
cannot be dimmed unless the armament
system circuit breakers are on and the
MASTER ARM switch is in either STBY
or ARM TEST SWITCH. The TEST
position illuminates the entire caution
light system. Testing the system will not
change any malfunction indication
existing prior to testing. There is no
provision for the copilot to reset his
MASTER CAUTION light.

Fire Warning System.

Fire warning lights are located in the FIRE 1
PULL and FIRE 2 PULL handles on the pilot
instrument panel (figure 1-24). Two lights are
located in the copilot/gunner glare shield and
indicate FIRE ENG 1 and FIRE ENG 2 when
illuminated. The pilot and copilot/gunner lights
are connected in parallel and both sets of lights
illuminate when energized. A FIRE WARN TEST
switch is located on the pilot instrument panel.
The switch is spring-loaded to the off position. The
TEST position causes all four fire warning lights
to illuminate aviation red, indicating the system is
operational. Excessive heat in engine 1

1-43

Section I
Part 2

NAVAIR 01-H1AAB-1

compartment causes FIRE 1 PULL and FIRE
ENG 1 lights to illuminate. Excessive heat in
engine 2 compartment causes FIRE 2 PULL and
FIRE ENG 2 lights to illuminate. Power is
supplied by the 28 vdc essential bus and protected
by ENG NO. 1 FIRE DET and ENG NO. 2 FIRE
DET circuit breakers.

CAUTION

* imHHHHHHHHHV '

Do not actuate the FIRE WARN TEST
switch more than 15 seconds. Prolonged
use will overheat the detector elements.

FIRE EXTINGUISHER SYSTEM OPERATION.

Illumination of either fire warning light indicates
excessive heat in the respective engine compart¬
ment (figure 1-25). Pulling the FIRE PULL
handle will shut off fuel to the affected engine, de¬
activate the ECU and rain removal circuits, close
the particle separator door, and arm both fire
extinguisher bottles. Positioning the FIRE EXT
switch to either MAIN or RESERVE will discharge
selected bottle into the affected engine compart¬
ment (Figure 1-25). To use the remaining bottle,
move the FIRE EXT switch to the opposite
position. Fire light illumination is not required to
discharge the extinguishers. Power is supplied by
the 28 vdc essential bus and protected by two
circuit breakers marked FIRE EXT MAIN and
FIRE EXT RSV.

NOTE

Pulling a FIRE PULL handle and
positioning the FIRE EXT switch to
MAIN will result in that bottle being
discharged into the selected engine
compartment. If both FIRE PULL
handles are pulled out and the FIRE
EXT switch is moved to MAIN position,
the bottle will not discharge. If the
switch is moved to the RESERVE
position, only one bottle will discharge.

The discharged extinguishing agent will
be routed to both engine areas and could
be ineffective in either engine area.

Fire Extinguisher.

A portable fire extinguisher is located on the
bulkhead to the left of the copilot/gunner seat.

First Aid Kit.

An aeronautical type first aid kit is located on the
aft bulkhead of the pilot compartment.

Survival Kit.

(Space Provisions only)

CREW COMPARTMENT DOORS.

Pilot and copilot/gunner access is provided by
canopy doors that are hinged at the top and swing
outward and up. The pilot canopy door opening is
on the right side and the copilot/gunner canopy
door is on the left side. Both doors are opened or
closed either electrically or manually from inside
or outside. Both doors may be stopped electrically
in any position between open and one inch from
closed. Both doors have switches located near the
door handles for electrical operation (figure 1-26).
To open either door manually from the closed
position, turn door handle and raise door to desired
position and trip (up) manual clutch release (figure
1-26). To close either door manually or to open door
from a position other than closed, depress (down)
manual clutch release and move the door. Clutch
may be engaged manually or electrically. Key
locks are provided in each door handle.

CANOPY JETTISON SYSTEM.

A canopy jettison system provides for rapid crew
egress in emergency situations (figure 1-26). The
system consists of a linear explosive system, used
to cut the side windows from the canopy support
structure, three canopy jettison handles, and the
interconnecting lines of flexible confined
detonating cord. ARM/FIRE mechanisms are
manually activated percussion type detonators.
When fired, all four window cutting assemblies
will be immediately detonated to blow the four
side windows outward in fragments, leaving empty
frames for exit or access. Jettison handles are
located on the pilot instrument glareshield,
copilot/gunner right console, and in the nose for
rescue crew. The system can be actuated by any
one of the three handles. The canopy jettison
handles have safety pins to prevent accidental
firing of the system. The pins must be pulled
before the system can be actuated. To actuate
the system, rotate any handle 90 degrees counter¬
clockwise and pull. Stowage for safety pins is
provided.

1-44

NAVAIR 01-H1AAB-1

Section I
Part 2

FIRE

| RPM

| MASTER 1

FIRE

ENG 1 |

| CAUTION|

ENG 2

FIRE

EXT

ROTOR 1
BRAKE |

I MASTER II
| CAUTION ||

RPM

FIRE

DET MAIN

FIRE

EXT

FiRE

DET RSV

-LIGHTS-1

O00

ROTOR RPM
CAUTION BK WRN

FIRE 1 PU LL 1 | FIRE 2 PULL

NOMENCLATURE

FUNCTION

ROTOR BRAKE
'MASTER CAUTION
'RPM
'FIRE 1-2

Circuit breakers

FIRE WARN TEST switch

Illuminate when rotor brake is pressurized.

Illuminate to alert crew to check caution panel.

Illuminate when rotor rpm is high or low or when
gas producer is low.

Illuminate to show fire in respective engine
compartment.

Protect respective electrical circuit.

Test fire warning lights.

*Repeater lights on copilot/gunner instrument panel.

210900-120

Figure 1-24. Warning/Caution Lights

Section I
Part 2

NAVAIR 01-H1AAB-1

NOMENCLATURE

FUNCTION

FIRE EXT switch

FIRE 1/2 PULL handle

MAIN — Fires main fire bottle.

OFF — Deactivates fire bottles.

RESERVE — Fires reserve fire bottle.

Illuminate to show fire in respective engine
When pulled;

Shut off fuel to respective engine.

Deactivate ECU and rain removal circuits.
Close particle separator door.

Arm both fire extinguisher bottles.

Circuit breakers

Protect electrical circuits

FIRE WARN TEST switch

Test fire warning lights.

210900-119

Figure 1-25. Fire Extinguishing System

1-46

NAVAIR 01 -HI AAB-1

Section I
Part 2

NOMENCLATURE
CANOPY JETTISON HANDLE
DOOR ACTUATOR
DOOR HANDLE
DOOR SWITCHES
DOOR LOCK

FUNCTION

Removes glass from doors and windows.

Position door.

Locks door in closed position and deactivates electrical circuit.

Opens or closes door electrically.

Secure helicopter doors.

210900-118

Figure 1-26. Entrance/Egress Systems

1-47

Section I
Part 2

NAVAIR 01 -HI AAB-1

Helmet visors shall be down prior to
activation of the CRS to preclude
possible eye injury.

PILOT SEAT.

The pilot seat is a vertically adjustable,
nonreclining type, installed at a reclined angle of
15 degrees (figure 1-3). The vertical height
adjustment is on the left side of the seat. The back,
bottom and side panels are made of ceramic and
fiberglass composite armor.

Additional protection is provided by side shoulder
panels which can be installed on or removed from
the basic seat. The seat is equipped with seat and
back cushions. A lap safety belt and inertia-reel
shoulder harness is also installed.

COPILOT/GUNNER SEAT.

The copilot/gunner seat is a fixed seat, installed at
a reclined angle of 15 degrees (figure 1-4). The seat
is made of ceramic and fiberglass composite
armor. The seat is equipped with a lap belt, inertia-
reel shoulder harness, plus seat and back
cushions. Arm rests are provided for each side of
the seat.

SHOULDER HARNESS.

An inertia-reel and shoulder harness is installed
on the pilot and copilot/gunner seats with a
manual lock-unlock control handle. With the
control in the unlocked position, the reel cable will
extend to allow the occupant to lean forward.
However, the reel will automatically lock when the
helicopter encounters an impact force in excess of
two G deceleration. Locking of the reel can be
accomplished with the harness at any position,
and the reel will automatically take up the slack in
the harness. To release the lock it is necessary to
lean back slightly to release tension on the lock
and move the control handle to the lock and then
unlock position. It is possible to have pressure
against the seat whereby no additional movement
can be accomplished and the lock cannot be
released. If this condition occurs, it will be

necessary to loosen shoulder harness. The reel
should be manually locked for takeoff and
landing.

VENTILATING SYSTEM.

Ventilating air is supplied through the air inlet
located on the leading edge of the pylon fairing.
Outside air is routed through an electrical blower
into the distribution system (figure 1-27). The
ECU/VENT switch is mounted on the pilot
instrument panel (figure 1-28). Placing the
ECU/VENT switch to VENT, actuates the
system. The pilot has adjustable outlets on his
instrument panel and controllable outlets on each
side of the instrument panel. The outlets on the
instrument panels shroud provide air for
defogging the canopy area. Controllable
ventilating air is also routed through the pilot seat
and back cushion. The copilot/gunner has one
instrument panel mounted adjustable outlet.
Ventilating air is also routed through the
copilot/gunner seat and back cushion. Air volume
through the instrument panel outlet is regulated
by the butterfly in the outlet. Power is supplied by
the 28 vdc essential bus and protected by the
VENT BLO circuit breaker.

Ventilating System Operation.

The ventilation system serves as a backup to the
copilot/gunner air conditioning system. This
system provides crew ventilation through two
pilot outlets, and two copilot/gunner outlets.

Defrosting/Defogging.

The ECU provides heated air for defrosting.
Heated air is directed to the side areas of the
canopy.

Rain and Ice Removal System.

Removal of rain or ice from the forward window
panel is accomplished by placing RAIN RMV
switch to RAIN RMV (figure 1-29). The switch is
located in the ECU panel on the pilot right console.
When the RAIN RMV switch is actuated, bleed air
valves open and bleed air mixed with outside air is
directed to the base of the forward windshield.

1-48

NAVAIR 01 -HI AAB-1

Section I
Part 2

Figure 1 -27. Environmental Control System Schematic

1-49

Section I NAVAIR 01 -HI AAB-1

Part 2

NOMENCLATURE
ECU/OFF/VENT switch

COOL WARM knob

FUNCTION

ECU — supply conditioned air to the crew
compartment.

OFF — deactivate system.

VENT — supply ambient air to the crew
compartment.

Adjust temperature of conditioned air.

Seat air knob

Adjust volume of air to the respective seat.

Circuit breaker

Protect electrical circuit.

DEFROST PULL

Adjust volume of air to the canopy.

210900-131

Figure 1-28. Environmental Control System

1-50

NAVAIR 01 -HI AAB-1

Section I
Part 2

NOMENCLATURE

FUNCTION

RAIN RMV/OFF switch
Rain removal nozzle

RAIN RMV — supply bleed air to the rain removal nozzle
OFF — remove power from system
Distribute air over windshield.

210900-130

Figure 1-29. Rain Removal System

1-51

NAVAIR 01-H1AAB-1

Section I
Part 2

Environmental Control Unit (ECU)

The ECU is located in the hydraulic compartment.
Conditioned air from the ECU is supplied through
the air distribution system. The ECU/VENT
switch, in the ECU position, actuates the unit. The
COOL/WARM knob regulates the temperature of
the outlet air. The COOL/WARM knob is located
on the control panel in the pilot right console.
Power is supplied by the 28 vdc essential bus and
protected by the ECU PWR circuit breaker (figure
1-29. The rain RMV and ECU switches should be in
the OFF or VENT position during flight conditions
requiring maximum engine performance due to
reduction in engine power available.

Power is supplied by the 28 vdc essential bus and
protected by the ECU PWR circuit breaker
(figure 1-28).

EXTERIOR LIGHTS.

Navigation Lights.

The navigation lights are controlled from the
LIGHTS control panel. Two switches are provided
for control of the navigation lights, FLASH-OFF-
STEADY and BRT-DIM. Power is supplied by the
28 vdc essential bus and protected by the NAV
LIGHTS circuit breaker.

Anti-collision Light.

The anti-collision light is mounted on top of the
engine cowl (figure 1-30). The light is controlled by
a switch in the LIGHTS control panel. The switch
is a two-position switch marked ANTI-COLL LT
ON-OFF. Electrical power for the light is supplied
from the 28 vdc essential bus and protected by the
ANTICOLLISION LIGHTS circuit breaker.

Fuselage Formation Lights.

The fuselage formation lights consist of five green
lights (figure 1-30). One is located on top of the 90
degree gearbox, two on top of the pylon fairing,
and one on the top surface of each wing tip. The
lights are controlled from the LIGHTS control
panel. A FUSELAGE switch is provided for
turning the formation lights on and varying the
brightness from off to bright. Power is supplied by
the 115 vac nonessential bus and protected by the
FORM LT PWR and FUS circuit breaker.

Rotor Tip Formation Lights.

The rotor tip formation lights are two white lights,
one installed on each rotor tip (figure 1-30). The
lights are controlled from the LIGHTS control
panel. A ROTOR TIP switch is provided for
turning the lights on and varying the brightness
from OFF to BRT. Power is supplied by the 115 vac
nonessential bus and protected by the ROTOR
FORM LT circuit breaker.

Searchlight.

The controllable searchlight is located in the
bottom fuselage section beneath the
copilot/gunner station. Control switches are
provided for both pilot and copilot/gunner. The
pilot control switches are located in the collective
stick switch box (figure 1-31). The switches are
marked SRCH LT - EXT/RET/L/R and SRCH
LT ON/OFF/STOW.

The copilot/gunner switches are located in the
miscellaneous panel. The switches are marked
SRCH LT, ON-OFF AND EXT-RETR. The pilot
has the capability to rotate the light right or left,
the copilot/gunner does not have this capability.
Power is supplied by the 28 vdc essential bus and
protected by the SRCH PWR and SRCH CONT
LIGHTS circuit breakers.

INTERIOR LIGHTS.

Crew Compartment Lights.

The pilot and copilot/gunner cockpit lights are
located on the side armored seat panels (figure 1-
31). Rheostat operating switches for the lights are
mounted on the light assembly body. Brightness is
controlled by operation of the rheostat. Rotation of
the lens clockwise provides white lighting,
counterclockwise red lighting. The rheostat is also
the ON-OFF switch for the light assembly. Power
is supplied by the 28 vdc essential bus and
protected by the CKPT LIGHTS circuit breaker.

Pilot, Copilot/Gunner Instrument Lights.

The pilot, co-pilot/gunner instruments contain
internal lighting. A pilot rheostat marked INST
LTS is located on the LIGHTS panel (figure 1-31).
A copilot/gunner rheostat marked INST LT is
located on the miscellaneous control panel. Power
is supplied by three 5 vdc instrument lighting
power supplies. One for the copilot and two for
the pilot; these are in turn powered by the 28

1-52

NAVAIR 01-H1AAB-1

Section I
Part 2

210900-129

1. FORMATION LIGHTS

2. AFT NAVIGATION LIGHT (RH NOT SHOWN)

3. ANTICOLLISION LIGHT

4. LEFT NAVIGATION LIGHT (RH NOT SHOWN)

5. SEARCHLIGHT (NOT SHOWN)

Figure 1 -30. Exterior Lighting

Section I NAVAIR 01-H1AAB-1

Part 2

Kr' ftt !- /5=;

000

ANTI-

CKPT COLLISION NAV

SRCH SRCH
PWR CONT

000

PLT PLT PLT &
INSTR INSTR GUNNER

SRCH LT switch

SRCH LT spring switch

Circuit breakers
INST LTS knob

ON — activates search light.

OFF — deactivates search light.

STOW — stows search light.

EXT — extends search light to vertical position.
R — turns light to the right.

L — turns light to the left.

RET — retracts light to horizontal position.
Protects electrical circuit.

OFF — deactivates lights.

BRT — adjusts light to full intensity.

210900-128

Figure 1-31. Interior Lights (Sheet 1 of 2)

1-54

NAVAIR 01-H1AAB-1

Section I
Part 2

NOMENCLATURE

FUNCTION

CSL LTS

ANTI-COLL LT

NAVIGATION LTS switches
FLASH/OFF/STEADY

BRT/DIM

FORMATION LIGHTS knobs
FUSELAGE

ROTOR TIP

MAP LIGHT
NOTE

Copilot/gunner compartment
not shown.

INSTR LT knob

CSL LT knob

SRCH LT switches
ON/OFF

EXT/RET

Map Light

OFF — deactivates lights.

BRT — adjusts light to full intensity.

ON — activates light.

OFF — deactivates light.

FLASH — activates lights flashing cycle.
OFF — deactivates lights.

STEADY — activates lights.

Controls brilliance.

OFF — deactivates lights.

1 to BRT — controls brilliance.

OFF — deactivates lights
1 to BRT — controls brilliance.

Provides pilot with moveable lighting.

OFF — deactivates light.

BRT — adjusts light to full intensity.

OFF — deactivates light.

BRT — adjusts light to full intensity.

ON — activates light.

OFF — deactivates light.

EXT — extends search light to vertical position.
RETR — retracts light to horizontal position.
Provides copilot/gunner with movable lighting.

210900-9-2A

Figure 1-31. Interior Lights (Sheet 2 of 2)

1-55

Section I
Part 2 - Part 3

NAVAIR 01 -HI AAB-1

vdc essential bus. Circuit protection is provided
by the PLT INST LTS and GUNNER INSTR LT
circuit breakers.

Pilot, Copilot/Gunner Console Lights.

The pilot, copilot/gunner consoles contain
internal lighting. A rheostat marked CSL LTS is

located on the pilot LIGHTS panel (figure 1-31),
and copilot/gunner miscellaneous control panel.

Power is supplied by the 28 vdc essential bus and
protected by the CSL PLT and GUNNER LIGHTS
circuit breaker.

PART 3 — SERVICE AND HANDLING

FUELING AND SERVICING.

Servicing points are presented on the Servicing
Diagram (figure 1-32). See figure 1-33 for
specifications. See figure 1-34 for system
capacities. See figure 1-35 for turning radius and
ground clearance, figure 1-36 for rotor blade
danger area.

Crew and Truck.

Only authorized and qualified personnel shall
operate fueling equipment. The plane captain
shall be responsible for fueling the helicopter after
each flight. He will make a visual check to ensure
the proper fuel is used. Do not locate the helicopter
in the vicinity of possible sources of ignition, such
as blasting, drilling or welding operations. A
minimum of 50 feet should be maintained from
other aircraft and 75 feet from any operating radar
set. Aircraft servicing vehicles will be positioned
parallel to the helicopter during any servicing
operation.

Grounding.

Prior to fueling, grounding devices on helicopter
and on trucks shall be inspected by fueling
personnel for proper ground.

Electrical Hazard.

Turn off all switches and electrical equipment in
helicopter. Check that no electrical apparatus,
supplied by outside power (electrical cords, drop
lights, floodlights, etc.) is in or near the helicopter.
For night fueling, safety flashlights shall be used.

Static pifferential.

Before using a fuel hose, the hose nozzle shall be
brought in contact with some metal part of the

helicopter, remote from the fuel cells to eliminate
any static differential that exists. This procedure
should result in eliminating static differential to
reduce the chance of static spark at fuel cell filler
port.

Attaching Wire Clamp.

Before removing the cell filler caps, the hose nozzle
ground attachment shall be connected to a metal
part of the helicopter at a safe distance from filler
openings and cell vents.

Fire Extinguishers and Attendant.

During fueling, a secondary operator or assistant
plane captain will man a C02 hand extinguisher
with a second extinguisher readily available.

ENGINE WASH PROCEDURES.

There are two types of engine wash procedures; the
engine performance recovery wash and the engine
desalinization rinse.

Only personnel designated in writing by
commanding officers shall be authorized to
conduct engine motoring wash procedures.

Engine Performance Recovery Wash

An engine performance recovery wash shall be
required when a deterioration in engine
performance is noted and/or the helicopter has
hovered below 30 feet over salt water.

Prior to the wash, the engine shall be allowed to
cool for a minimum of 40 minutes.

1. Armament — OFF/SAFE

2. FUEL - OFF

1-56

NAVAIR 01-H1AAB-1

Section I
Part 3

1. COMBINING GEARBOX OIL FILLER

2. ENGINE OIL FILLER INSIDE COWLING

3. ENGINE OIL SIGHT GAGE WINDOW

4. NO. 1 HYDRAULIC SIGHT GAGE WINDOW

5. ENGINE OIL SIGHT GAGE LIGHT SWITCH

6. WING TANK FILLER - SAME ON RIGHT SIDE

7. TAIL ROTOR GEARBOX FILLER

8. TAIL ROTOR GEARBOX SIGHT GAGE

9. INTERMEDIATE GEARBOX SIGHT GAGE

10. INTERMEDIATE GEARBOX FILLER INSIDE COWLING

11. ROTOR HUB RESERVOIRS

12. TRANSMISSION FILLER CAP

13. COMBINING GEARBOX SIGHT GAGE

14. NO. 1 HYDRAULIC FILTER WINDOW

15. ENGINE OIL FILLER INSIDE COWLING

16. TRANSMISSION SIGHT GAGE

17. ENGINE AND XMSN SIGHT GAGE LIGHT SWITCH

18. NO. 2 HYDRAULIC SIGHT GAGE AND FILTER

19. FUEL FILLER

20. GROUNDING RECEPTACLE

Figure 1-32. Servicing Diagram

1-57

Section I
Part 3

NAVAIR 01-H1AAB-1

SYSTEM

SPECIFICATION

STANDARD

ALTERNATE

EMERGENCY

FUEL

JP-5

JP-4

JP-8(M 1L-T-83133)

ASTMD-1655

JETA/JetA-1

OIL

ENGINE SECTIONS

SHELL ASTRO 555*

COMBINING GEARBOX

SHELL ASTRO 555*

TRANSMISSION

SHELL ASTRO 555*

INTERMEDIATE GEARBOX

SHELL ASTRO 555*

TAIL ROTOR GEARBOX

SHELL ASTRO 555*

MAIN ROTOR HUB GRIPS

MIL-L-461 52

ANY HIGH DETER¬
GENT 10 W 30 OIL

HYDRAULIC

SYSTEMS

NO. 1 and NO. 2

MIL-H-83282

MIL-H-5606

NONE

*Preheating required for temperatures below-40°C.

N2/83

210900-44A

*NATO

SYMBOL

U S. MILITARY
SPEC MIL-J-5624
GRADES

U S. COMMERCIAL
SPEC ASTM D-1 655-62T
GRADES

U.K.

GRADES

F-44

***JP-5

None

AVCAT/48

F-40

JP-4

**JET B

AVTAG

F-34

None

*•

JET A-1

AVTUR/50

None

****JET A

None

* The NATO symbols denote general types of fuels as manufactured under several national
military and commercial specifications, and can be applied to products meeting a general category.
Fuels having the same NATO symbol are interchangeable for use by military aircraft.

** Equivalent to JP-4 except that freezing point is -60°F vice -76°F.

*** F-44 approved fuel afloat — F-44 and F-40 approved fuel ashore.

**** Equivalent to JET A-l except freezing point is -38°C vice -50°C.

Figure 1-33. Specification Sheet

1-58

Change 1

NAVAIR 01 -HI AAB-1

Section I
Part 3

SYSTEM

CAPACITIES

AVAILABLE

UNUSABLE

FUEL

FORWARD FUEL CELL

190 US GAL

AFT FUEL CELL

123 US GAL

313 US GAL

2 US GAL

RIGHT WING TANK

100 US GAL

0.42 US GAL

LEFT WING TANK

100 US GAL

1.15 US GAL

513 US GAL

3.57 US GAL

OIL

TOTAL

ENGINE SECTION 1

3 US QUARTS

3.4 QUARTS

6.4 QUARTS

ENGINE SECTION 2

3 US QUARTS

3.4 QUARTS

6.4 QUARTS

COMBINING GEARBOX

1 US QUART

4 QUARTS

5 QUARTS

TRANSMISSION

15-1/2 QTS

18-1/2 QUARTS

INTERMEDIATE GEARBOX

3-1/2 PINTS

TAIL ROTOR GEARBOX

4-1/2 PINTS

MAIN ROTOR HUB GRIP No. 1

2 US QUARTS

2 QUARTS

MAIN ROTOR HUB GRIP No. 2

2 US QUARTS

2 QUARTS

HYDRAULIC

SYSTEM No. 1

9 PINTS

SYSTEM No. 2

11 PINTS

♦TRANSMISSION OIL COOLER Cl
CONTAINS AN ADDITIONAL 3 Q
WHICH IS TO BE CONSIDERED |
UNUSABLE.

RCUIT

lUARTS

Figure 1 -34. System Capacities

1-59

Section I
Part 3

NAVAIR 01-H1AAB-1

NOTE

MINIMUM GROUND CLEARANCES

9 FT. 0 IN.

2 FT. 10.3 IN.
4 FT. 6 IN.

1 FT. 8.4 IN.*
1 FT. 3 IN.

♦CHECK ANTENNAS THAT MAY PROTRUDE LOWER

Main Rotor Blades Stationary
Tail Skid

Tail Rotor Blades Stationary
Bottom of Fuselage
Bottom of Turret

210900-126

Figure 1-35. Turning Radius on Ground Handling Wheels

1-60

NAVAIR 01 -HI AAB-1

Section I
Part 3

Figure 1 -36. Rotor Blade Danger Area (Sheet 1 of 2)

1-61

Section I
Part 3

NAVAIR 01-H1AAB-1

Figure 1-36. Rotor Blade Danger Area (Sheet 2 of 2)

3. BATTERY - ON

4. APU - CONNECTED (If available)

5. Throttles — OFF

6. START switch — ON (for engine being
washed) 30 second limit.

7. START switch — OFF (one minute off)

8. Repeat steps 6 and 7 for second engine.

Allow the cleaning solution to soak for a
minumum of 10 minutes and maximum of 30
minutes, then rinse twice with fresh water
utilizing steps 6 and 7.

The following dry cycle is optional except for cases
where antifreeze has been added to fluids:

1. START switch — ON (for engine being dried)
30 second limit.

2. START switch — OFF.

3. Repeat steps 1 and 2 for second engine.

Engine Desalinization Rinse

An engine desalinization rinse is required after the
last flight of the day when deployed aboard ship,
and/or when operating from bases within two
miles of salt water or flown below 500 feet over salt

1-62

NAVAIR 01 -HI AAB-1

Section I
Part 3

water. Prior to the wash, the engine shall be
allowed to cool for a minimum of 40 minutes.

1. Armament — OFF/SAFE

2. FUEL - OFF

3. BATTERY - ON

4. APU — CONNECTED (If available)

5. Throttles — OFF

6. START switch — ON (for engine being
rinsed) 30 second limit,

7. START switch - OFF

8. Repeat steps 6 and 7 for second engine.

9. Repeat steps 6 and 7 for first engine.

10. Repeat steps 6 and 7 for second engine.

The following dry cycle is optional except for cases
when antifreeze has been added to fluids.

CAUTION

* **»*»++%+»**»*»**vv*v» >

Allow starter to cool for 5 minutes.

1. START switch — ON (for engine being dried)
30 second limit.

2. START switch — OFF

3. Repeat steps 1 and 2 for second engine.

PRESSURE HOT FUELING

Pressure hot fueling is prohibited when
ordnance is onboard except during
operational contingency.

1. Throttles - ENG IDLE.

2. Copilot/gunner — OUT (as required).

3. Copilot/gunner door — CLOSED.

4. Pilot door — CLOSED.

5. Helmet visor — DOWN.

6. TANK INTCON - OPEN.

7. CROSS FEED - OPEN.

8. FORCE TRIM - ON.

Fueling Personnel.

1. Helicopter - GROUND.

2. Fueling unit — GROUND.

3. Fire guard — POST.

4. Filler cap — REMOVE.

5. Fuel probe - CONNECT TO RECEIVER.

6. Fuel handle — ON.

7. Push and hold one of two PRECHECK
REFUEL plungers on rim of fueling valve.
Fuel flow should stop. Release plunger to
allow flow. Repeat with other PRECHECK
REFUEL plunger.

8. Fuel vent - CHECK FOR AIR FLOW.

CAUTION

<»»»»»»»»»»»»»+»»»»+»»» >

Pressure fueling shall be discontinued
immediately when any of the following
is observed:

• Fuel flow does not stop on
either PRECHECK
REFUEL check.

• No air flow out of vent.

• Slow or no fuel flow.

• Fuel out of vent.

• Fuel seeping out around
vent access panel.

• Fuel quantity gage shows
no increase.

• Sound of structural
deformation.

9. After satisfactory tests, continue fueling until
automatic shutoff when cells are full.

1-63

NAVAIR 01-H1AAB-1

Section I
Part 3

10. Fuel handle — OFF.

11. Fuel probe - DISCONNECT.

12. Filler cap — INSTALL.

13. Grounds — REMOVE.

\ -1

CAUTION

If fuel quantity gage shows below full
quantity when pressure fuel shutoff
occurs, a malfunction or failure exists
within the fuel system.

Ground Crew Requirements.

Normal helicopter fueling crew.

Emergency Shut Down.

PILOT

1. Throttles — Close.

2. FIRE PULL - PULL (Both Handles).

3. BATTERY — OFF.

4. EXIT HELICOPTER.

PRESSURE FUELING.

1. BATTERY switch — ON.

2. TANK INTCON switch - OPEN.

3. BATTERY switch — OFF.

4. Helicopter — GROUND.

5. Fueling unit — GROUND.

6. Fire guard — POST.

7. Filler cap — REMOVE.

8. Fuel probe - CONNECT TO RECEIVER.

9. Fuel handle — ON.

10. Push and hold one of two PRECHECK
REFUEL plungers on rim of fueling valve.
Fuel flow should stop. Release plunger to
allow flow. Repeat with other PRECHECK
REFUEL plunger.

11. Fuel vent - CHECK FOR AIR FLOW.

CAUTION

Pressure refueling shall be discontinued
immediately when any of the following
is observed:

• Fuel flow does not stop on
either PRECHECK
REFUEL check.

• No air flow out of vent.

• Slow or no fuel flow.

• Fuel out of vent.

• Fuel seeping out around
vent access panel.

• Fuel quantity gage shows
no increase.

• Sound of structural
deformation.

12. After satisfactory tests, continue fueling until
automatic shutoff when cells are full.

13. Fuel handle — OFF.

14. Fuel probe - DISCONNECT.

15. Filler cap — INSTALL.

16. Grounds — REMOVE.

CAUTION

If fuel quantity gage shows below full
quantity when pressure fueling shutoff
occurs, a malfunction or failure exists
within the fuel system.

LINE OPERATIONS.

The primary function of the flight line section is to
ensure the safest and most efficient operation of all
ground level activities including the elimination of
foreign object damage (FOD). The Flight Line

1-64

NAVAIR 01 -HI AAB-1

Section I
Part 3

Officer is charged with supervision of all such
activities and is guided in this mission by OPNAV
INST. 4790.2.

Limitations for Towing the Helicopter.

CAUTION

;;

Towing the helicopter on the ground
handling gear (ground handling
wheels), on unprepared surfaces, at high
gross weights may cause permanent set
in the aft crosstube.

TOWING SPEED.

Towing speed will not exceed five miles per hour.
Sudden stops and starts shall be avoided. Extreme
caution shall be exercised when towing in a
congested area.

GROUND HANDLING GEAR TYPES.

Two types of ground handling gear can be used for
moving the helicopter, forward mounted and aft
mounted.

GROUND HANDLING GEAR. At gross weights of
13,560 pounds or less,the ground handling gear
may be used for moving the helicopter. Early
models do not have hand brakes. While in move¬
ment, each wheel assembly should be manned
by a qualified aircraft handler. A qualified
aircraft handler shall be positioned on the
tail skid to take the weight off the front of the
skid tube and to provide steerage. Two aircraft
handlers may be utilized on the tail skid when
wind/weight conditions warrant. The helicopter
may be towed or pushed by hand if a sufficient
number of aircraft handlers are available. Care
should be exercised when lowering the helicopter
onto the deck. The helicopter should be lowered
slowly and assure all personnel are well clear.

WING WALKER.

When towing a helicopter near hangars, obstruc¬
tions, or other aircraft, a wing walker, equipped
with a whistle, shall be stationed on each side of
the helicopter to ensure adequate clearance. At
night the wing walker will carry a flashlight or
luminous wand and the helicopter position lights
shall be turned on.

MOVEMENT.

During all phases of helicopter movement the
main rotor blades shall be secured to the tailboom
by means of a rotor tiedown.

Operation of Equipment:

Only qualified personnel shall operate towing
equipment. Towing couplings shall be inspected
prior to towing. Only approved tow bars shall be
used. Ground handling wheels shall be installed in
eye bolts provided on each landing gear skid tube,
located forward of aft cross tube and forward of the
forward cross tube. Reference maintenance manual
for proper ground handling gear installation and
operation.

Care shall be taken to ensure that the
ground handling pins are properly
installed into eye bolts on the skid tube.

FORWARD MOUNTED GROUND HANDLING
GEAR. The forward ground handling gear should
be used when helicopter is at a high gross weight
and/or forward of mid eg.

PROPER OPERATION WHEN FORWARD
MOUNTED GROUND HANDLING GEAR IS
USED. Install all ground handling gears in eye
bolts on skid tube.

Extend aft ground handling wheel on one side
only. Extend forward gear on same side. Extend
remaining aft ground handling gear. Extend
remaining forward mounted ground handling
gear. Lower in reverse order.

Do not raise or lower forward mounted
ground handling gear unless the aft
ground handling gear is raised.

One hand brake is installed on each aft ground
handling gear assembly. 1 )uring actual movement
of the helicopter each hand brake shall be manned
by a qualified aircraft handler.

A qualified aircraft handler shall be positioned on
the tail skid to provide steerage. The helicopter
may be towed or pushed by hand. Care should be
exercised when lowering the helicopter onto the
skids. The helicopter should be lowered slowly
after assuring all personnel are well clear.

1-65

Section I
Part 4

NAVAIR 01-H1AAB-1

PART 4 — OPERATING LIMITATIONS

If aircraft rotor or engine limitations are
exceeded, record on the VIDS/MAF.
Further flight shall not be attempted
until aircraft is inspected by qualified
maintenance personnel.

INSTRUMENT MARKINGS.

Refer to figure 1-37 for instrument markings.

TORQUE LIMITS (ENGINE).

1. Refer to figure 1-37.

2. Duration of single engine operation at or
above 49.4% torque shall be entered on the
yellow sheet.

STARTER LIMITATIONS.

The duty cycle for the starter in this installation is
as follows:

1. 30 seconds on — 1 minute off.

2. 30 seconds on — 1 minute off.

3. 30 seconds on — 30 minutes off.

After 30 minutes the duty cycle can start over
again. *

ROTOR BRAKE LIMITATIONS.

1. Do not apply rotor brake above 60% Nr.

2. Do not apply rotor brake below 25% Nr.

3. Do not move handle above detent while rotor
is turning.

4. For rotor brake start: Release by 61%Ng.

AIRSPEED LIMITS.

Decrease airspeed 5 KIAS for each 1000 feet of
density altitude above 4000 feet. Airspeed limits
below 4000 feet density altitude are as follows:

1. 190 KIAS without stores.

2. 170 KIAS any configuration with stores.

3. 120 KIAS steady state autorotation.

4. Sideward flight 35 knots.

5. Rearward flight 30 knots.

6. Airspeed indicator is unreliable at airspeeds
less than 40 knots.

7. Flight within the red area of height velocity
diagram should be avoided (figure 1-38).

8. Canopy doors shall not be opened in flight
at airspeeds in excess of 45 KIAS.

9. Auxiliary fuel tanks. Data not available at
this time.

PROHIBITED MANEUVERS.

1. No acrobatic maneuvers permitted (acrobatic
as defined in OPNAVINST 3710.7 series).

2. Flight below + 0.5 G is prohibited.

3. No practice full autorotations unless gross
weight is 12,500 or less, and a qualified
instructor, designated for full autorotations
by the commanding officer, is in the cockpit.

4. Practice autorotation entries within the
shaded areas of the height velocity diagram
(figure 1-38) are prohibited.

5. No airstarts or manual fuel control operation
above 15,000 feet.

6. No dual engine throttle chops above Vh
(maximum level) flight speed attainable at
MC power.

7. No solo flight permitted from the
copilot/gunner cockpit.

MINIMUM CREW REQUIREMENTS.

NOTE

The view is restricted from the aft
cockpit. With only one pilot in the
helicopter a slight sideslip may be
required to see the landing area during
final approach.

1-66 Change 1

NAVAIR 01-H1AAB-1

Section I
Part 4

ENGINE TORQUE

( 1 0—45.2% — Normal Operating Range

I 145.2—53.1% — 30 Minute Limit
■^53. 1—55.6% — 10 Second Limit

TRANSMISSION TORQUE

Dive — 65% Maximum for Any Airspeed
Above The Maximum Airspeed
Obtainable at 84.9% Torque

Maximum Continuous Power — 84.9% Up to
Level Flight Maximum Speed
Maximum Power — 85—100% for 5 Minute Limit

|100% Maximum

ROTOR TACHOMETER (Nr)
POWER ON:

IHH91 % — Transient Minimum
I 1 97—100% — Normal Operation
98—100% — Dives
005% — Maximum

105 — 120% — Transient Maximum
(10 Seconds)

POWER OFF:

84% — Transient Minimum
91—105% — Normal Operation
105 — 120% — Transient Maximum
(10 Seconds)

ENGINE TACHOMETER (Nf)

POWER ON:

^ 1 */,^

? 5:

..do i'* ' 30 =

V>90 RPM

K, 80 -50^?

70 60 vN

1 1 11

GAS PRODUCER TACHOMETER (Ng)

I 1 97 - 100% — Normal Operation

98 - 100% — Dives
■■102% — Maximum

110%— Transient Maximum
(10 Seconds)

102 - 110% - Transient Maximum
(10 Seconds) (Engine
logbook entry required
anytime 102% Nf is
exceeded)

] 102% Ng - Maximum

102 — 103% — Transient Maximum (10 Seconds)
(Engine logbook entry is required
anytime 102% is exceeded.

N2/83

210075-153-1 B

Figure 1 -37. Operating Limits and Instrument Markings
(Sheet 1 of 3)

Change 1

1 67

Section I
Part 4

NAVAIR 01-H1AAB-1

ENGINE OIL TEMPERATURE
HHl16°C Maximum

TRANSMISSION OIL TEMPERATURE
Hi 10°C Maximum

ENGINE OIL PRESSURE

40 PSI — Minimum
80-115 PSI — Normal Operation
115 PSI — Maximum

TRANSMISSION OIL PRESSURE

BBB 30 PSI — Minimum
f 1 40-70 PSI — Normal Operation
H70 PSI — Maximum

AC-DC VOLTMETER

Generator ON — 27 TO 28.5 Volts DC
AC Voltage — 113.5 +5 Volts

DUAL AMMETER
I 1 0—150 Normal Operation

COMBINING GEARBOX OIL TEMPERATURE

Jl 1 6°C Maximum

COMBINING GEARBOX OIL PRESSURE

40 PSI Minimum
60—85 PSI Normal Operation
85 PSI Maximum

FUEL PRESSURE

5 PSI — Minimum
5-25 PSI — Normal Operation
25 PSI — Maximum

210075-153-2B

1-68

Figure 1-37. Operating Limits and Instrument Markings
(Sheet 2 of 3)

NAVAIR 01-H1AAB-1

Section I
Part 4

INTER TURBINE TEMPERATURE
INFLIGHT:

200—789°C — Maximum Continuous Power
837°C — Maximum 30 Minute Limit
837°C — Light Illumination
838— 900°C — Transient Acceleration
Maximum 5 Seconds.

STARTING:

838 — 900°C — 5 Second Limit
901—1150°C — 2 Second Limit
Over 1150°C — Overtemp Inspection

DUAL HYDRAULIC PRESSURE

2940—3060 PSI Normal Operation
] 2200—3200 PSI Operation Range
Limits are for gage pressure variations with
control or SCAS inputs. Steady LOW
or HIGH pressure readings are an indication
of hydraulic system malfunction.

13200 Maximum

21007 5-153-3A

Figure 1-37. Operating Limits and Instrument Markings
(Sheet 3 of 3)

1-69

Section I
Part 4

NAVAIR 01-H1AAB-1

DATA BASIS: ESTIMATED

CONDITION:

ALL CONFIGURATIONS
CALM WIND

INDICATED AIRSPEED - KIAS

Figure 1-38. Height Velocity Diagram

210900-79

1-70

NAVAIR 01-H1AAB-1

Section I
Part 4

The minimum crew requirements for the AH-1T
(TOW) helicopter consists of a pilot.

CENTER OF GRAVITY LIMITATIONS.

Refer to figure 1-39 for center of gravity
limitations.

LATERAL CG LIMITATIONS.

Most right or most left lateral eg limits is 6.0
inches. Within this limit the helicopter may be
flown with a single store on any station. Stores on
both stations on the same side, with opposite side
empty, can possibly exceed the lateral limit
(depending on the particular stores).

NOTE

The most critical flight regime with the
lateral eg at the most right station is a
level flight at full power. The most
critical flight regime with the lateral eg
at the most left station is a 120 KIAS
autorotation. If the lateral eg limits are
exceeded there may not be sufficient
lateral control margin to maintain
balanced flight.

ACCELERATION G LIMITATIONS.

At 10,000 pounds gross weight: 0.5 to 3.5 G. At
14,000 pounds: 0.5 to 2.5 G (figure 1-40).

Refer to figure 1-40 for G limits.

FLIGHT WITH CROSSTUBE FAIRINGS
REMOVED.

Crosstube fairings are optional equipment. Flight
with fairings removed is authorized to the limits
of the basic aircraft. Flying qualities, performance,
and structural characteristics are unchanged.

DUMMY TSU FERRY FLIGHTS.

Flight with the dummy TSU installed is authorized
for ferry flights only to the following maximum
limits:

1. Airspeed —120 kt.

2. Bank Angle — 30 deg.

3. Acceleration — Plus 1.0 to plus 1.5 G.

Change 1 1-71

GROSS WEIGHT - POUNDS

Section I
Part 4

NAVAIR 01-H1AAB-1

192 194 196 198 200 202

FUSELAGE STATION - INCHES

NTSA

210900-42

Figure 1-39. Center of Gravity Diagram

1-72

LOAD FACTOR

NAVAIR 01-H1AAB-1

Section I
Part 4

AH-1T (TOW AND NON TOW)

210900-153

Figure 1*40. Gross Weight Vs Acceleration Nz

1-73/(1-74 blank)

NAVAIR 01-H1AAB-1

Section II

SECTION II — INDOCTRINATION

TABLE OF CONTENTS

Introduction.2-1

Ground Training Syllabus .2-1

Pilot Ground Training.2-1

Pilot Flight Training.2-1

Flight Crew Designation,

Qualifications and

Requirements.2-2

INTRODUCTION.

The operating procedures contained in this
manual will apply to AH-1T (TOW) helicopters
when performing assigned missions within their
capabilities. The information contained herein is
to clarify, amplify, and standardize those areas
where there is room for variance of interpretation
by individual commands. The procedures
contained herein cannot possibly cover every
conceivable situation, but are intended to govern
situations most frequently encountered. The
safety and success of any mission are of
paramount importance with precedence of actions
depending upon the existing situation.

GROUND TRAINING SYLLABUS.

A ground training program shall be established
which will ensure thorough training and a high
degree of readiness for all flight personnel. The
ground training syllabus which follows is to be
used as a guide and represents the minimum
requirements to be met prior to completing the
familiarization stage in the flight training
syllabus as set forth by the type commands.

PILOT GROUND TRAINING.

1. Every pilot checking out in the AH-IT (TOW)
helicopter will be required to complete a
course of instruction in the AH-1T (TOW).
This course of instruction will vary from
about 20 hours to the maximum of 40 hours
depending upon the pilot's background.

2. A written examination will be given on the
NATOPS Flight Manual and NWP series
publications.

Qualifications.

Crew Requirements.

Currency .

Personal Flying Equipment

3. Instruction and examination must be
completed on the following subjects prior to
completion of the flight familiarization
phase:

a. Helicopter operational performance
(flight characteristics, systems operation,
etc.)

b. Weight and balance.

c. Publications (FAA, Tactical, Technical
and associated, etc.)

d. Communications.

e. Survival and first aid.

f. Search and rescue.

g. Flight planning, fuel management.

h. Helicopter navigation.

i. Flight safety.

j. Emergency procedures.

PILOT FLIGHT TRAINING.

A flight training syllabus shall be established by
each command to accomplish maximum training
for the mission and tasks assigned. The syllabus
must be flexible and tailored to fit the situation
and the varying nature of the tasks and
commitments. The flight training syllabus will
contain the following phases: familiarization,
formation, instruments, navigation, night,
shipboard, and special categories.

2-1

Section II

NAVAIR 01-H1AAB-1

FLIGHT CREW DESIGNATION,
QUALIFICATIONS AND REQUIREMENTS.

The flight crew qualifications and requirements as
set forth in the following paragraphs are minimums
and are not to be interpreted as limiting in any way
the establishment of higher requirements by proper
authority.

Designation.

A naval aviator or aviation pilot will be designated
as qualified in model only after he has previously
been designated as a helicopter pilot under the
provisions of OPNAVINST 3710.7 series. A pilot
who has qualified in one of the helicopter
classifications shall have a certificate thereof,
signed by the qualifying authority. This certificate
will state the model helicopter and modification
thereto in which he is qualified and shall be placed
in his Aviator’s Flight Log Book, Officers
Qualification Jacket, or Enlisted Service Record
Book, as appropriate.

Designating Authority.

Commanding officers, or higher authority in the
chain of command are empowered to designate pilot’s
qualified in model and issue certification thereto.
The immediate superior in command to the
commanding officer, or higher authority, may
assume the function. The authority assuming the
function shall issue appropriate instructions.

Qualifications.

PILOT QUALIFIED IN MODEL (PQM).

A pilot qualified in model must have satisfactorily
completed a Flight Training Syllabus or
demonstrate comparable proficiency to include the
capability of executing all assigned missions and
tasks, and must further meet the requirements as
set forth in detail in OPNAVINST 3710.7 series,
and completed satisfactorily a NATOPS
evaluation.

MISSION COMMANDER.

A mission commander shall be a properly
qualified Naval aviator designated by appropriate
authority. The mission commander may exercise
command over single Naval aircraft or formations

of Naval aircraft. He shall be responsible for all
phases of the assigned mission except those
aspects of safety of flight which fall under the
perogatives of individual pilots in command.
Requirements for designation as mission
commander will be outlined by appropriate
authority.

SECTION LEADER.

A section leader must be a pilot qualified in model. In
addition, this pilot must be fully qualified to lead a
section under all conditions in performance of any of
the squadron tasks.

DIVISION LEADER.

A division leader must be a pilot qualified in model
with no less than 600 total flight hours. Of this total,
200 hours must be in helicopters of which 50 hours
must be in squadron model.

FLIGHT LEADER.

A flight leader must be a qualified division leader
with no less than 750 total flight hours.
Consideration will be given to rank and experience,
when warranted, to allow for exceptions by the
commanding officer.

COPILOT/GUNNER.

A copilot/gunner is a pilot who has completed the
FAM stage and all FRONT COCKPIT ORDNANCE
requirements.

FUNCTIONAL CHECK PILOT.

A FCP must have a minimum of 100 hours in model
PQM and be designated in writing by the unit
commanding officer.

CREW REQUIREMENTS.

1. A pilot designated as qualified in model shall
command the helicopter and occupy one of the
control positions on all service and combat
flights.

2. A transition pilot (pilot under instruction),
rated safe for solo, may command the helicopter

2 2

NAVAIR 01-H1AAB-1

Section II

on all types of operational training missions
within his capabilities and which, in the
opinion of the commanding officer, is best suited
to instill pilot confidence and helicopter
command responsibilities.

3. On all flights, a qualified observer may
occupy the forward seat to ensure adequate
visual surveillance. A qualified observer is
anyone who is thoroughly briefed in cockpit
conduct and safety, to include intercom
system operation and lookout
responsibilities.

4. All instructional flights will be under the
direct supervision of a designated PQM.
Familiarization stage training will be
conducted only under VFR conditions.

Currency.

ANNUAL FLYING AND CURRENCY
REQUIREMENTS.

To ensure that the skill of naval aviators is
maintained at an acceptable standard of readiness
for fleet operations, the annual flying
requirements as set forth in OPNAVINST 3710.7
series must be adhered to by all active duty naval
aviators.

CREW REST REQUIREMENTS.

Pilots should not be scheduled for more than
6-1/2 hours of normal flying per day. Eight hours
for 1 day is permissible provided a minimum of 2
hours crew rest is taken between each 4 hour
period of flight. A basic crew day of 12 hours from
first brief to last shutdown shall not be exceeded.
Minimum ground time after extended flight
operations shall be sufficient time for crew
members to eat and obtain at least 8 hours of
uninterrupted rest. Exceeding the above crew
rest requirements may result in crew fatigue
causing impaired judgement and reduced
performance.

NATOPS EVALUATION.

On assignment to another unit a PQM will not be
required to receive a NATOPS evaluation if the log
book entry and pilot’s qualification jacket indicates
successful completion of the check within the last 12
months.

WAIVERS.

Unit commanders are authorized to waive in writing
minimum flight and/or training requirements where
recent experience in similar model helicopters
warrants.

PERSONAL FLYING EQUIPMENT.

The latest available type of flight safety and survival
equipment listed below shall be worn by all pilots and
crewmembers on all flights unless a tactical combat
environment or military exigency require on site
deviations. See OPNAVINST 3710.7 series for
further details.

All Flights.

1. Protective helmet.

2. Flight safety boots/field shoes.

3. Nomex gloves.

4. Nomex flight suit.

5. Identification tags.

6. Survival knife and sheath.

7. Personal survival kit.

8. Signalling device for all night flights and for all
flights over water or sparsely populated area.

9. Parachutes will be utilized in accordance with
OPNAVINST 3710.7.

Over Water Flights.

1. Life preservers shall be worn.

2. Anti-exposure suits shall be provided for all
personnel in accordance with OPNAVINST
3710.7 series.

Night and Instrument Flights.

1. A flashlight shall be carried in the helicopter.

2. Approach plates.

3. Maps.

Change 1

2-3/(2-4 blank)

NAVAIR 01 -HIAAB-1

Section III
Part 1

SECTION III — NORMAL PROCEDURES

TABLE OF CONTENTS

PART 1 - FLIGHT PREPARATION

Mission Planning.3-1

PART 2 - SHORE-BASED PROCEDURES

Introduction.3-5

Scheduling.3-5

Ground Operations.3-6

Discrepancy Reporting.3-6

Exterior Inspection.3-6

Pre-Entry Inspection.3-6

Interior Inspection — Copilot/Gunner.3-7

Interior Inspection — Pilot. 3-8

Pre-Start Checklist. 3-9

Start Checklist.3-9

Post Start Checklist.3-11

Pre-Takeoff Checklist.3-11

Air Taxiing.3-12

Types of Takeoff.3-12

After Takeoff.3-13

Climb.3-13

Cruise.3-13

Descent.3-13

Pre-Landing Check.3-13

Landing.3-14

Autorotation Practice.3-16

Hovering Autorotation.3-16A

Dual Engine Failure (Simulated).3-16A

Quick Stop.3-16A

Twenty and Thirty Degree Dives.3-16A

Practice High Speed Low Level
Autorotations.3-16B

Manual Fuel Flight.3-16B

Tail Rotor Malfunction.3-16C

Dual Hydraulic Failure (Simulated).3-16C

Waveoff.3-16C

Shutdown.3-16D

Postflight External Inspection.3-16D

Night Flying.3-17

PART 3 - SHIP-BASED PROCEDURES

Command Responsibility.3-17

Field Carrier Landing Practice.3-17

Carrier Qualification.3-18

Operation of Equipment.3-19

Flight Deck Operations.3-19

Air Capable Ship Operations.3-27

Night Operations.3-28

Debriefing.3-28

PART 4 - SPECIAL PROCEDURES

Full Autorotation Landing.3-29

Formation and Tactics.3-29

Rendezvous.3-33

Formation Takeoffs and Landings.3-33

PART 5 — FUNCTIONAL CHECKFLIGHT
PROCEDURES

Introduction.3-35

Requirements.3-35

Procedures.3-35

Profile.3-36

PART 1 — FLIGHT PREPARATION

MISSION PLANNING.

Introduction.

Adequate and thorough planning of the flight is
necessary to assure the successful completion of
any mission.

Factors Affecting Helicopter Lift Capability.

TEMPERATURE.

High free air temperatures (fat) result in increased
inlet air temperatures which have an adverse
effect on the power output of gas turbine engines.

On the T400-WV-402 engine, one percent loss in
horsepower can be expected for each 1 degree
celsuis above standard day temperature. For each
2 degrees Celsius above standard, approximately
0.1 percent decrease in GAS PROD (Ng) rpm for
maximum power can be expected. An increase in
(fat) at a constant pressure altitude causes an
increase in density altitude, which results in
decreased hover performance.

HUMIDITY.

The effect of humidity on gas turbine engines is
negligible.

3-1

Section III
Part 1

NAVAIR 01 -HIAAB-1

ALTITUDE.

Altitude has a marked effect on the performance of
all aircraft engines. Air density and temperature
decrease as altitude increases. As air density
decreases, the mass flow of air through the gas
turbine decreases. However, the gas turbine
operates more efficiently at the lower temperatures
0 encountered at high altitudes. At altitude, the
power output of gas turbine engines decreases as
evidenced in the cockpit by a decrease in the torque
pressure reading and the specific fuel consumption
(engine fuel consumption in pounds per hour
divided by engine shaft horsepower) decreases due
to increased engine efficiency. With the collective
pitch control set, the ENG RPM (Nf) will begin to
droop as higher altitudes are reached. Operating
rpm can be reestablished by reducing the angle of
attack of the blades (by decreasing collective).

WIND.

If a helicopter can takeoff and land into a steady
wind, its payload can be increased because less
power is required for the same flight performance
with wind than without wind. Helicopters
operating from the decks of ships underway are in
an excellent position to take advantage of the
relative wind generated by the ship movement.
However, an allowance for deck edge and elevator
turbulence must be made. Consideration must be
given to winds in the landing zone ashore when at
maximum gross weight conditions.

GROUND EFFECT.

For hovering flight closer than one-half rotor
diameter to the earth, the lifting ability of a
helicopter is increased by ground effect. Since the
power required to hover increases with an increase
in height above the ground, the helicopter can
hover at heavier gross weights in-ground effect
(IGE) than out-of-ground effect (OGE).

Weight Limitations Applicable to Helicopters.

THE AERODYNAMIC — POWER WEIGHT
LIMIT.

Increases in ambient air temperature and/or
altitude restrict lift capability of the helicopter
because a decrease in air density will result in
decreased power available from the engine and a
loss of rotor efficiency. The relationship of lift
capability to atmospheric conditions is found in
the performance charts in section XI. While flight
operations based on HIGE limit will permit an
increase of lift capability, HOGE weight
computations will be used for normal training
operations. Exceptions to this will be necessary
when operational and service flights are made
under favorable conditions which require carrying
payloads at an altitude beyond the capability of
the helicopter to HOGE. Sliding landings and
takeoffs will further increase payloads but require
a surface of sufficient length in an area free of
obstacles. HOGE and HIGE should be computed
prior to takeoff or landing. Operations based on
these exceptions should be made only under
carefully calculated requirements.

WEIGHT AND BALANCE.

The AH-1T (TOW) is a class IB aircraft for weight
and balance purposes (the CG limits can be
exceeded by some normal loads) and therefore
needs loading control. The Manual for Weight
and Balance, NA01-1B-40, includes guidance and
data for the specific serial number aircraft to
insure proper loading control.

The maximum allowable gross weight for takeoff
is 14,000 pounds and must not be exceeded. This
weight is determined by structural and flying
qualities flight tests done by NAVAIR and Bell.
CG limits are shown in figure 1-39. The lateral
CG limit is 6.0 inches right or left.

Form 365F of NA01-1B-40 is not normally
required for each flight if a current form is on file.
See OPNAVINST 3710.7 series for further infor¬
mation. The AHC will ensure that the maximum
allowable gross weight, longitudinal and lateral
limits will not be exceeded during flight.

3-2

NAVAIR 01-H1AAB-1

Section III
Parti

General Precautions.

Special care will be exercised to avoid flying over
populated areas, civilian airports, turkey and
chicken farms, etc. In all cases conformance with
existing regulations is mandatory.

Requirements for Mission Planning.

Mission planning has two requirements. The first
requirement is for pilot and operations personnel
to calculate normal and emergency helicopter
operating capabilities concurrent with existing
ambient conditions and mission requirements
prior to every flight on a daily basis. The second
requirement is preparation of planning documents
for a future helicopter assault or support mission
and is normally prepared from weather
summaries and predicted weather in the area to be
considered. Weather summaries suitable for
preparation of such estimates can be prepared or
obtained by any authorized weather facility with a
forecasting capability. Fuel reserve for all flights
shall be computed so as to land with no less than
10% or 20 minutes, whichever is greater.

COMPUTATION CARD.

The computation card for determining capabilities
(figure 3-1) shall be used for mission planning.
Deviations and substitutions may be made within
the standard form. Substituting HIGE or HOGE is
an example. HOGE computations should be made
of the mission requirements in order to have this
information readily available during flight.

BRIEFING.

The pilot is responsible for briefing the crew. This
briefing shall ensure complete understanding of
the mission. The pilot shall give specific
instructions to cover special situations that may
occur.

A briefing guide will be used. On training flights,
the appropriate syllabus guide should be used.
Each pilot will maintain a kneepad and record all
flight numbers, call signs and other data
necessary to successfully assume the lead and
complete the assigned mission. The briefing guide
will include the following items.

GENERAL.

1. Helicopter call signs, event.

2. Lead/alternates.

3. Fuel load, stores, gross weight.

4. Start, taxi, takeoff times.

5. Takeoff data, rendezvous.

TARGET OR DESTINATION.

1. Primary.

2. Secondary.

3. Operating area, targets.

4. Control agency.

5. Time on station or over target.

NAVIGATION/FLIGHT PLANNING

1. Duty runway.

2. Climbout.

3. Operating/restricted areas.

4. Obstacles to flight.

5. Mission plan.

6. Cockpit coordination.

7. Bingo/low fuel.

8. Holding.

9. Approach/lighting.

10. GCA/missed approach.

11. Recovery.

12. Divert/emergency fields.

COMMUNICATIONS

1. Frequencies.

2. Agencies.

Change 1

3-3

Section III NAVAIR 01-H1AAB-1

Part

WIND

KNOTS

1 .

TAKEOFF PRESSURE ALT FT. FAT °C.«

NO-WIND HOGE MAX GRWT

WIND CORRECTION

(+)

HOGE MAX GRWT/WIND

OPERATING WEIGHT (BASIC WT., CREW, MISC.)

(-)

PAYLOAD PLUS FUEL

RANGE OUT NMI, TIME/FUEL OUT MIN.

FUEL FOR START, TO, AND RESERVE

TOTAL FUEL REQUIRED

10.

TOTAL FUEL ABOARD (GAGE READING)

(-)

11.

PAYLOAD OUT LB

12.

LANDING PRESSURE ALT. FT., LAND

TEMP °C WIND

13.

LANDING NO-WIND HOGE MAX GRWT

14.

WIND CORRECTION (LANDING)

(+)

15.

LANDING HOGE MAX GRWT/WIND

16.

FUEL OUT

(+)

17.

*MAX GRWT PERMITTED AT TO DUE

TO LANDING CONDITIONS

*lf this weight is less than line (4) above, substitute this weight
in line (4) and recompute payload.

NOTE: This form may be abbreviated for daily local area operation.

Figure 3-1. Computation Card

3-4

NAVAIR 01 -HI AAB-1

Section III
Part 1 — Part 2

3. Procedure/discipline.

3. Nav aid failures.

4. IFF.

4. Loss of visual contact/VMC/IMC.

5. Nav Aids.

5. Inadvertent IMC.

6. Signals.

WEAPONS

1. Loading.

2. Arming.

3. Hot ordnance routes.

6. Lost procedures.

7. SAR.

o. Helicopter/system failures.
9. Crew coordination.

SPECIAL INSTRUCTIONS

4. Pattern.

5. Switches.

6. Airspeeds.

7. Minimums.

8. G versus weight.

9. Duds, hung ordnance, de-arm, jettison.

10. Safety.

fj 11. Crew coordination.

WEATHER

1. Local, enroute, destination/forecast.

2. Alternates.

3. Winds.

EMERGENCIES

1. Aborts.

2. Radio failures.

1. Intelligence.

2. Safety.

3. Reports/authentication.

DEFRIEFING.

A proper debriefing conducted under tactical or
training conditions can be the most important part
of a flight. Mistakes can be discussed in an
atmosphere free from distractions. Under tactical
conditions debriefing is a primary source of
information leading to the location of targets,
distribution of troops, and many other important
considerations. An outline should be followed
when debriefing a flight. This outline should
contain all of the items for briefing plus the
following:

1. All unusual circumstances encountered.

2. Discrepancies arising.

3. Constructive criticism can be conducted in
such a manner that all concerned can
participate and present their ideas on the
conduct of the flight.

PART 2 — SHORE-BASED PROCEDURES

INTRODUCTION.

Shore-based procedures are discussed in this
chapter to cover as many operational situations as
possible.

SCHEDULING.

The commanding officer, or his designated
representative, is responsible for the promulgation
of the flight schedule when based ashore. The

Change 1 3-5

Section III
Part 2

NAVAIR 01 -HIAAB-1

flight schedule, when published, becomes an order
of the commanding officer. The flight schedule will
contain sufficient information to assure all
preparations relative to flight can be
accomplished in a smooth and timely manner. The
minimum essential items which shall be included
on the flight schedule are found in OPNAV INST
3710.7.

GROUND OPERATIONS.

Preflight Inspection.

Prior to flight, the pilot and aircrewmen shall
conduct a complete visual check of the helicopter.

Fire Guard.

Prior to starting the engine, a qualified fire guard
shall be stationed near the engine and remain in
readiness with a fire bottle until the engine is
operating.

• The fire guard shall remain clear of the
exhaust and compressor blade area.

• Ear protection and goggles that provide
adequate peripheral vision shall be worn
by flight line personnel.

helicopter Acceptance.

Die pilot in command shall ensure, prior to accept¬
ance, that the helicopter is satisfactory for safe
flight and can accomplish the assigned mission. The
pilot will review the VIDS/MAF discrepancies from
the last ten flights and all previous outstanding
discrepancies. The pilot shall conduct a thorough
preflight inspection.

1. The pilot shall ensure that the plane captain
has conducted a standard daily preflight as
set forth in the NAVAIR 01-H1AAB-6 series
and signed the yellow sheet prior to each
flight.

2. The pilot, when satisfied with the yellow
sheet information, will sign applicable
portions of the yellow sheet.

DISCREPANCY REPORTING.

Immediately following each flight, the pilot shall
note all discrepancies in detail by completing the

applicable items of the VIDS/MAF form in
accordance with OPNAVINST 4790.2 series. To
aid in descrepancy analysis, specific information
such as position of controls, movement of
controls and results, instrument readings, etc.,
should be recorded in flight, if practical, to be
included on the yellow sheet. Maintenance
troubleshooters should be available for consul¬
tation. The pilot will ensure that he has conveyed
his complete knowledge of the discrepancy orally
and in writing.

EXTERIOR INSPECTION.

On pilot’s first two FAM hops, exterior preflight
inspection shall be conducted in accordance with
MRC cards. Figure 3-2 represents minimum
preflight inspection for all flights.

PRE-ENTRY INSPECTION.

1. Armament cb switches — UP/INBD.

2. Armament switches — OFF/SAFE.

3. FIRE handles —IN.

4. FIRE extinguishers — OFF.

5. BATTERY - OFF.

6. INVERTER — MAIN.

7. FUEL switches — ON.

8. SCAS - ON.

9. ANTI COL LT - ON.

10. FUEL QTY — CHECK.

11. SCAS - OFF.

12. FUEL switches — OFF.

13. INVERTER - OFF.

14. BATTERY — OFF.

15. Rotor tiedowns — REMOVE/STOW.

16. Engine covers — REMOVE/STOW.

If single pilot operation is to be conducted, the
following items shall be checked in the copilot/
gunner cockpit:

1. Safety bolt and shoulder harness —
SECURE.

2. Loose equipment — SECURE.

3. Canopy jettison handle pin — IN.

4. UHF EMER — OFF, COVERED.

3-6 Change 1

NAVAiR 01 -HIAAB-1

Section III
Part 2

1 Windshield
Searchlight.

2. Pilot canopy for proper operation and latching.
Hydraulic fluid level and leaks.

Fuel quantity and cap security.

Right side skid tube and cross tubes.

Right wing.

Transmission oil level.

Engine No. 2 oil level.

Combining gearbox oil level.

Oil or fuel leaks in transmission and engine
compartment right side.

All access doors and panels secured right side.
Avionics and battery compartment.

Main rotor blades from right side.

Elevator for damage right side.

4. 42 degree gearbox oil level and leaks.

90 degree gearbox oil level.

Elevator for damage left side.

Fuselage and tailboom left side.

Avionics compartment.

Oil or fuel leaks in transmission and engine
compartment left side.

Fire bottles.

Engine No. 1 oil level.

Hydraulic fluid level.

Rotor mast and head.

Pitot tube cover removed.

All access doors and panels left side secure.

Left skid tube and cross tubes.

Left wing.

5. Gunner canopy for proper operation and latching

3. Tail rotor.

Tail skid.

Main rotor blade tiedown removed.

6. Turret gun drive assembly.

Rear gun mount.
Declutcher-feeder cannon plug.
Elevation and azimuth brakes.
Turret cowling.

Muzzle and mid-barrel clamps.
Telescopic sight unit.

210947-2

Figure 3-2. Exterior Inspection

5. INTER panel — SET.

6. ENG 1 GOV, ENG 2 GOV — AUTO.

7. INSTR LT and CSL LT — OFF.

8. SRCH LT — OFF.

9. F TRIM - ON.

10. ELEC PWR — ELEC PWR.

11. WG ST JTSN — OFF, COVERED.

12. PILOT OVERRIDE switch — OFF.

13. Armament switches and controls — AS
REQUIRED.

14. Canopy door — CLOSE AND SECURE.

INTERIOR INSPECTION-COPILOT/
GUNNER.

These items represent minimum inspection for all
flights.

1. Canopy door — AS DESIRED.

2. Loose equipment — STOW AND SECURE.

3-7

Section III
Part 2

IMAVAIR 01 -HIAAB-1

3. Pedals — ADJUST.

4. Safety belt and shoulder harness —
FASTEN.

0 5. Inertial reel — CHECKED (including self¬

locking feature).

6. Canopy jettison handle pin — IN.

7. PILOT OVERRIDE - OFF.

8. Armament switches and controls — AS
REQUIRED.

9. UHF EMER - OFF, COVERED.

10. INTER panel - SET.

11. ENG 1 GOV, ENG 2 GOV - AUTO.

12. INSTR LT and CSL LT - AS REQUIRED.

13. F TRIM - ON.

14. SRCH LT - OFF.

15. ELEC PWR - ELEC PWR.

16. WG ST JTSN - OFF, COVERED.

17. Inform pilot — COPILOT/GUNNER
CHECKLIST COMPLETE.

INTERIOR INSPECTION - PILOT.

1. Rotor tiedowns — REMOVED.

2. Seat and pedals — ADJUST.

3. Shoulder harness — ADJUST.

B 4. Inertial reel — CHECKED (including self¬
locking feature).

5. Canopy door — AS DESIRED.

6. Right circuit breakers — IN.

7. Lights - AS DESIRED (ANTI-COL - ON).

8. PITOT HTR - OFF.

9. RAIN RMV - OFF.

' 10. COMPASS - SLAVED.

11. IFF— OFF.

12. TACAN — OFF.

13. ADF — OFF.

14. KY-28 —OFF.

15. UHF — OFF.

16. FM-OFF.

17. ALE-39 ARM switch - SAFE.

18. ALE-39 PWR switch — OFF.

19. APR-39 — OFF.

20. Clock - SET AND RUNNING.

21. RADAR ALTITUDE altimeter — OFF.

22. ALT — SET.

23. FIRE EXT - OFF.

24. Canopy jettison handle pin — IN.

25. CODE HOLD — OFF.

26. FIRE PULL handles - IN.

27. ECU/VENT - OFF.

28. MASTER ARM — OFF.

29. KY-28 — OFF.

30. RADAR BEACON - OFF.

31. ALQ-144 — OFF.

32. EMERGENCY JETTISON SELECT-OFF.

33. INTER panel - AS DESIRED.

34. SCAS POWER - OFF.

35. SMOKE ARM - OFF.

36. ENG 1, ENG 2 FUEL - OFF.

37. TANK INTCON - OPEN.

38. CROSS FEED — OPEN.

39. ENGINE 1, ENGINE 2 GOV - AUTO.

40. ENGINE 1, ENGINE 2PART SEP-AUTO.

41. NO. 1 GEN, NO. 2 GEN - OFF.

42. INVERTERS - OFF.

43. HYD - ON.

3-8 Change 1

NAVAIR 01 -HI AAB-1

Section III
Part 2

44. NON-ESS BUS — NORMAL.

45. FORCE TRIM — ON.

46. BATTERY - OFF.

47. AUX FUEL - OFF.

48. Left circuit breakers — IN, TOGGLE
SWITCHES INBOARD.

49. SRCHLT - OFF.

50. Collective strap — OFF.

51. Copilot/gunner checklist — COMPLETE.

INTERIOR INSPECTION (NIGHT FLIGHTS).
In addition to interior inspection for all flights the
pilot shall inspect the following:

1. Flashlights — AVAILABLE.

2. All interior lights — CHECK OPERATION.

3. All exterior lights - CHECK OPERATION.

PRE-START CHECKLIST.

1. Helmet — ON.

2. BATTERY — ON, WITH OR WITHOUT
APU (22 vdc minimum for start).

I^ARNIN^I

If battery voltage is below 22v, replace
the battery before preceeding with start
up. A low voltage battery will cause
battery damage, and possible damage to
the aircraft and/or injury to personnel.

3. APU (if required) — CONNECT (check
voltage 26 minimum, 29 maximum).

4. INVERTERS — MAIN (AC Voltage — 113.5
± 5).

5. Rotor brake — ON (if desired) check light.

6. FIRE WARN — TEST.

NOTE

Do not hold test switch on for more than
15 seconds.

7. FUEL GA TEST - PRESS.

8. MASTER CAUTION, CAUTION-
ADVISORY panel — TEST and RESET.

A qualified pilot shall be in the pilot seat whenever
the engine and rotor are started. Prior to start the
rotor tiedowns shall be removed and the
surrounding area clear of unnecessary personnel,
equipment and obstructions. The pilot shall
acknowledge plane captain/fire guard all clear
signals prior to start. Helicopter tiedowns shall be
removed with caution when the engines and rotors
are operating and only upon proper signal.

CAUTION

Tail winds in excess of 10 knots with
external power or battery start may
result in smoke, tailpipe fire, and/or
excessive ITT.

To help equalize engine starter wear, start the
number one engine on odd calendar days and the
number two engine on even calendar days.

9. Throttles - OPEN, CHECK IDLE STOP,
OFF.

10. Throttle friction second engine to be started
-SET.

11. RPM — DECR (5 seconds).

12. FUEL first start engine — ON (Pressure 5-25
PSI, FUEL BOOST lights out).

13. Main rotor — CLEAR.

14. VOLTS DC - MINIMUM 22 VOLTS.

START CHECKLIST.

If for any reason a starting attempt is
discontinued, allow the engine to come to a
complete stop and then accomplish a 15 second
motoring run. Repeat the complete starting
sequence.

If INLET TEMP fails to rise within 10 seconds
after opening the throttle, close the throttle,
START OFF, and FUEL OFF. Allow the fuel to
drain for one minute and then accomplish a 15
second motoring run. If battery voltage will not
rise above 15.5 vdc after engaging starter, turn the
starter off and utilize an external power source. Do
not move the START switch through the off
position and engage the other starter.

If rotor slipping occurs after the starter is engaged
and the decision is made to continue the start, the
rotor brake shall be released immediately. If rotor
slipping occurs in winds greater than 35 knots, the
start shall be aborted.

Change 1 3-9

Section III
Part 2

NAVAIR 01-H1AAB-1

1. START first engine — ON.

9. Waming/CAUTION lights - CHECK.

a. Observe a positive indication of oil
pressure.

b. When GAS PROD (Ng) passes 12%, check
battery voltage; if above 15.5 vdc smoothly
roll the throttle to the low side of the
IDLE STOP.

c. Monitor INLET TEMP to prevent engine
overtemperature.

CAUTION

Prepare for a hot start if ITT approaches
850 degrees C. For ITT above 837
degrees C, log temperature and time.

838 - 900 degrees C — 5 second limit.

901 - 1150 degrees C — 2 second limit.

over 1150 degrees C — Overtempera¬
ture inspection.

2. As GAS PROD (Ng) passes 50%, START —
OFF.

3. Rotor brake — RELEASE (below 61% GAS
PROD (Ng).

If severe main rotor flapping occurs due
to high/gusty winds, apply cyclic into
wind, as required, to prevent mast
bumping. If mast bumping occurs,
shut down helicopter.

4. Temperature/pressures — CHECK
(Combining gearbox, transmission and
engine).

5. APU - DISCONNECT AS REQUIRED.

6. GEN - ON, CHECK AMPERAGE.

7. IDLE STOP — RELEASE.

8. Throttle - INCREASE TO 85% ENG RPM
(Nf, SET FRICTION.)

NOTE

Do not operate engine in excess of 71%

Ng until engine and combining gearbox
temperature reach plus 15 degrees C.

10. Flight control - CHECK.

a. Check cylic movement — Displace cyclic
forward, aft, left and right approximately 2
inches. Note tip path plane correlation.

b. Directional pedals — Displace left and
right approximately 1 inch.

c. Collective — Raise collective sufficiently
to note ROTOR RPM (Nr) droop.

11. FORCE TRIM — CHECK.

a. With FORCE TRIM ON — Displace cyclic
and pedals; if operating properly, controls
will return to original position. Check that
gradient force is equal in all movements of
cyclic and pedals. Ensure that force trim
will hold controls in displaced position by
utilizing the cyclic stick force trim switch.

b. Turn FORCE TRIM OFF on pilot console.
No motoring of cyclic or pedals is allowed.

c. Displace cyclic and pedals to ensure force
gradient springs have released.

d. FORCE TRIM - ON.

12. Hydraulic check — Move the HYD switch to

theSYS. 1 OFF position,NO. 1HYDRPRESS

light is illuminated on the CAUTION
ADVISORY panel and HYD PSI1 near zero.
Tail rotor pedals will be stiff with collective
and cyclic normal. Make all movements slow
and small. Switch to theSYS. 2 OFF position,
NO. 2 HYDR PRESS light illuminates and
HYD PSI 2 near zero. HYD switch to ON and
all caution lights should be out, all controls
normal. MASTER CAUTION light will
illuminate each time a caution segment light
illuminates.

13. AMPS - CHECK (Below 150 amps).

14. FUEL second engine — ON.

15. Throttle friction second engine — OFF.

16. START second engine — ON.

a. Observe a positive indication of oil
pressure.

3-10 Change 1

NAVAIR 01 -HIAAB-1

b. When GAS PROD (Ng) stabilizes
(minimum 12%) smoothly roll the throttle
to the low side of the IDLE STOP.

c. Monitor INLET TEMP.

17. As GAS PROD (Ng) passes 50% START —
OFF.

18. IDLE STOP - RELEASE.

19. Throttle — Continue to increase to 85% ENG
RPM (Nf). Ensure that ENG RPM (Nf) does
not exceed first started engine. The torque of
the first engine should drop off slightly as the
ENG RPM (Nf) needles marry. A non-
engaged engine is indicated by ENG RPM
(Nf) slightly (2% or more) higher than the
engaged engine and a near zero torque
indication. If a non-engagement occurs,
smoothly close the throttle of the non-
engaged engine and when stopped, shut
down the engaged engine.

NOTE

If an abrupt ENG RPM (Nf)
deceleration, jolt, or noise occurs during
shutdown, do not attempt another start.

20. GEN - ON.

21. VOLTS DC - CHECK (27 - 28.5 vdc).

POST START CHECKLIST.

With both engines operating, maintain 85% ENG

RPM (Nf) while completing the checklist.

1. SCAS POWER — POWER.

2. Attitude gyro — SET.

3. RADAR ALTITUDE altimeter — ON and
SET.

4. IFF and radios — ON.

5. KY-28 - AS REQUIRED.

6. COMPASS — ALIGNED.

7. SCAS — CHECK NO-GO LIGHTS OUT.
(Engage channels and check release in both
cockpits.)

8. Throttles - SLOWLY FULL OPEN. CHECK
THROTTLE DECALS ALIGNED.

Section III
Part 2

9. RPM — 97 to 102 SET AT 100%.

10. ENG TRIM - SET.

11. Throttles - ENGINE IDLE - CHECK i
STOPS 62 ± 2%.

12. ROTOR RPM AUDIO - CHECK.

13. Throttles - FULL OPEN.

14. ROTOR RPM - AUDIO.

15. ECU/VENT - AS DESIRED.

CAUTION

RAIN RMV system should be turned
OFF as soon as cleared vision will
permit. Heat may melt windshield if
operated for a lengthy period on a dry
windshield. '

NOTE

A decrease in power available can be
expected when operating the ECU
and/or RAIN RMV.

PRE-TAKEOFF CHECKLIST.

1. RPM — 100%.

2. Caution and warning lights — CHECK.

3. TURRET STOW light - ON.

4. Instruments - CHECK PRESSURE AND
TEMPERATURES.

5. FUEL QTY - CHECK.

6. SCAS - ENGAGED.

7. ECU — OFF or VENT/RAIN RMV — OFF.

8. Shoulder harness — LOCKED.

9. MASTER ARM - OFF.

10. Area — CLEAR.

11. Hover power — NOTE POWER REQUIRED.

12. Canopy jettison pins — OUT (after panel
check and personnel are clear of helicopter).

3-11

NAVAIR 01 -HIAAB-1

Section III
Part 2

AIR TAXIING.

Movement of the helicopter from one ground
position to another can be accomplished by air
taxiing at an altitude of 3 to 5 feet (skid tube to
ground surface). From a hover, apply sufficient
cyclic to establish a slow rate of movement over the
ground in the desired direction. In confined areas,
this rate of movement should be no faster than a
man can walk.

Whenever possible, all air taxiing should be done
by pointing the nose of the helicopter in the desired
direction of movement. Sideward and rearward
flight may be necessary for use in high winds and
in confined areas. Due to increased rotor wash
caused by air taxiing, caution should be exercised
when in the vicinity of other aircraft due to rotor
turbulence. Particular attention is directed to the
increased rotor wash and its effects on loose
objects and debris in the vicinity of the helicopter.
Sufficient ground control personnel shall be
available to provide for the safe taxiing of
helicopters in the vicinity of obstructions or other
aircraft. Only approved standard taxi signals will
be used. Extreme caution should be exercised when
taxiing at night.

TYPES OF TAKEOFF.

Because of the versatility of helicopters and their
ability to takeoff from small areas, conditions at
the time of takeoff are the governing factors in the
type of takeoff to be accomplished. The factors
governing the type of takeoff to be accomplished
are the gross weight of the helicopter, the pressure
altitude, outside air temperature, prevailing
winds, the size of the takeoff area, and the tactical
situation. There are many possible variations in
takeoff procedures.

As the helicopter accelerates from hovering flight
to flight in any direction, it passes through a
translational period. If engine power, rpm, and
collective are held constant in calm air, a
momentary settling will be noted when the cyclic
control is moved forward to obtain forward speed.
This momentary settling condition is a result of
the helicopter moving from the ground cushion
and the tilting of the tip-path plane of rotation of
the main rotor blades to obtain forward airspeed.

Wind velocity at the time of takeoff will partially
eliminate this settling due to the increased airflow
over the main rotor blades. As wind velocity
increases this settling will be less pronounced.

Takeoff Performance.

A normal takeoff can be accomplished whenever
the helicopter is capable of hovering with the skids
5 to 10 feet above the ground. The hovering charts

in section XI can be used to determine if the
helicopter can hover out-of-ground effect and in-
ground effect.

Normal Takeoff to Hover.

The vertical takeoff is the normal type of takeoff,
and should be used whenever possible. The
helicopter is lifted from the ground vertically to a
height of approximately 3 to 5 feet where the flight
controls and engine may be checked for normal
operation before continuing to climb. A normal
vertical takeoff is made in the following manner:
Increase throttle to full open with the collective
pitch full down. Select desired rpm with the RPM
switch. Place cyclic control in the neutral position.
Increase collective pitch control slowly and
smoothly until hovering altitude of 3 to 5 feet is
reached. Apply pedals to maintain heading as
collective is increased. Make minor corrections
with cyclic to ensure vertical ascent, and use
pedals to maintain heading.

Normal Takeoff From Hover.

Hover briefly to determine if engine and flight
controls are operating properly. From a normal
hover at 3 to 5 feet altitude, apply forward cyclic to
accelerate into effective translational lift;
maintain hovering altitude with collective and
maintain heading with pedals, until translational
lift is attained and the ascent has begun. Just ■
prior to translational lift, the pilot will note a 0
slight decrease in Ng. In order to preclude sinking, Q
a slight increase in power may be necessary. As the I
aircraft is flown through translational lift (to n
preclude “ballooning”), a slight reduction in |
power may be necessary. Adjust power and I
smoothly lower the nose of the helicopter to B
arrive at approximately 25 feet of altitude and 50 j
knots of airspeed. Continue to accelerate and H
climb. Then smoothly lower nose of helicopter to w
an attitude that will result in an increase of air- |
speed to at least 7 0 knots. Adjust power as required to I
establish the desired rate of climb. Stabilize air-
speeed and torque as soon as smooth rate of
acceleration will permit.

Normal Takeoff From The Ground.

This takeoff is utilized for expeditious departure or
where normal takeoff to a hover is undesirable,
example; heavy sand or loose grass. With the
helicopter on the ground, coordinate increased
collective with simultaneous forward cyclic to
takeoff and move smoothly into translational lift.
Maintain normal takeoff attitudes until
translational lift is attained then proceed into
normal climb.

3-12 Change 1

NAVAIR 01-H1AAB-1

Section III
Part 2

Sliding Takeoff.

A sliding takeoff is made as follows: Apply power
until the helicopter is light on the skids. Smoothly
apply a slight amount of forward cyclic in order to
begin a ground slide. Do not attempt to rush the
forward motion of the helicopter as it will “dig in”.
As translational lift is gained, the helicopter will
fly off in a near level attitude. Maintain this
attitude until reaching 40 knots at which time a
normal climb may be initiated.

Maximum Power Takeoff.

Place cyclic in neutral position. With throttles full
open, increase collective smoothly. As the
helicopter leaves the ground, continue increasing
power to maximum available (not to exceed red
line) and assume a 80 knot attitude. As power is
increased, maintain heading by smoothly
coordinating pedals. When sufficient altitude for
obstacle clearance is attained, smoothly increase
airspeed and reduce power to establish a normal
climb.

Confined Area Takeoff.

This takeoff is utilized to depart an area over an
obstacle where little or no forward motion is
possible, until the helicopter is above the height of
the obstacle. Lift into a four-foot hover if possible
without exceeding limits. If within limits smoothly
increase collective to maximum allowable power
and lift straight up until skids are above obstacle
height. Apply forward cyclic and accelerate into
translational lift; then proceed into normal climb.

Crosswind Takeoff.

In the event a crosswind takeoff is required, there
will be a definite tendency to drift downwind. This
tendency can be corrected by applying the cyclic
into the wind a sufficient amount to prevent
downwind drift during takeoff. When a crosswind
takeoff is accomplished, it is advisable to turn the
helicopter into the wind for climb as soon as
obstacles are cleared and terrain permits.

AFTER TAKEOFF.

After the helicopter accelerates forward to 10 to 15
KIAS, less power is required to sustain flight due to
increase in aerodynamic efficiency as airspeed is
increased to best climbing speed. Takeoff power
should be maintained until a safe autorotative
airspeed is attained, then power may be adjusted
to*establish the desired rate of climb.

CLIMB.

The normal climb is made by adjusting nose attitude
to maintain at least 70 KIAS. Refer to section XI
for optimum climb airspeeds. At approximately
100 feet prior to the desired cruising altitude,
smoothly lower the nose, and allow the helicopter
to accelerate to cruise airspeed, while maintaining
a slight rate of climb to reach cruising altitude. As
the airspeed approaches cruise airspeed, adjust
power to maintain desired altitude and airspeed.

CRUISE.

Normal cruise will be conducted at a safe altitude
and as dictated by weather, helicopter
configuration and weight, terrain and obstacles,
mission of flight, safety of the helicopter, and
safety of persons and property on the ground.
Refer to section XI for design cruise airspeeds and
airspeed indicator corrections. Power and attitude
should be adjusted to attain desired cruise
airspeed.

DESCENT.

A descent is performed at a normal cruise airspeed
and collective pitch control as required for desired
rate of descent. At approximately 100 feet prior to
desired cruising altitude, adjust nose attitude and
power setting to level off at desired altitude.

PRE-LANDING CHECK.

1. RPM — 100%.

2. Caution and warning lights — CHECK.

3. TURRET STOW light — ON.

4. Instruments — CHECK PRESSURE AND
TEMPERATURES.

5. FUEL QTY — CHECK.

6. SCAS - ENGAGED.

7. ECU — OFF or VENT/RAIN RMV — OFF.

Change 1 3-13

Section III
Part 2

NAVAIR 01-H1AAB-1

8. Shoulder harness — LOCKED.

9. MASTER ARM — OFF.

| 10. Landing light - AS REQUIRED.

LANDING.

Normal Approach and Landing.

|j The downwind leg should be flown at 80 KIAS,

" 500 feet above the surface. Select the 180 degree
position with reference to the existing wind. At the
180 degree position commence a coordinated
descending turn, to arrive at the 90 degree position
at 300 feet and 70 KIAS. Adjust the rate of turn
and rate of descent so as to intercept the landing
line with 1,000 feet of straightaway and about 125
feet of altitude. At this point adjust nose altitude
smoothly to slow the airspeed and decrease the
rate of descent. Start the gas producer accelerating
by slightly increasing collective while the
helicopter is still in translational lift. Maintain
heading with pedals. The objective of a normal
approach is to simultaneously arrive over the
point of intended landing with zero ground speed
and approximately 3 to 5 feet of altitude. This
should be accomplished without an extreme flare
and/or abrupt power change.

Once the helicopter is established in a hover, lower
the collective to establish a slow, controlled rate of
descent to a gentle touchdown, making corrections
with pedals and cyclic to maintain a level attitude,
vertical descent, and constant heading. Upon
contact with the ground, continue to lower
collective smoothly 9nd steadily until the entire
weight of the helicopter is resting on the ground
and the collective is full down.

Slope Landing.

Make the slope landing by heading the helicopter
generally cross-slope. (Slope landing should be
made cross-slope with skid type gear.) Descend
slowly, placing the upslope skid on the ground
first. Coordinate reduction of collective pitch with
lateral cyclic (into the slope) until the downslope
skid touches the ground. Continue coordinating
reduction of collective and application of cyclic
into the slope until all the weight of the helicopter
is resting firmly on the slope. If the cyclic control
contacts the stop before the downslope skid is
resting firmly on the ground, return to hover, and
select a position where the degree of slope is not so

great. After completion of a slope landing, and
after determining that the helicopter will maintain
its position on the slope, place the cyclic in neutral
position.

After upslope skid contacts deck, a roll
rate must be established for the down-
slope skid to contact the deck. Angular
momentum can build to the point where
dynamic rollover can ensue regardless of
helicopter angle of bank. If mast bumping
occurs, reposition cyclic to stop bumping
and re-establish hover.

I h % hhhh%h%%%»»h% ,

CAUTION

If mast bumping occurs, reposition
cyclic toward center, keep control
inputs and aircraft roll rate small to
avoid dynamic rollover, then re-establish
hover.

Crosswind Landing.

Crosswind landings can generally be avoided in
helicopter operations. Occasionally, plowed,
furrowed or eroded fields, and narrow mountain
ridges may require that crosswind landings be
made. The crosswind landing in such instances is
utilized to prevent landing at a high tipping angle
or dangerous tail low attitude.

Crosswind landing may also be accomplished on
smooth terrain when deemed advisable by the
pilot. The following procedures should be observed
in accomplishing crosswind landing:

1. ENG RPM (Nf) 100 percent.

2. Hover helicopter crosswind.

3. Hold the cyclic control stick into the wind to
prevent side drift throughout the landing.

4. Proceed as in normal landing.

Steep Approach and Landing.

The steep approach procedure is a precision,
power-controlled approach used to clear obstacles

3-14 Change 1

NAVAIR 01

and to accomplish a landing in confined areas.
Slightly past the normal 180 degree position,
commence a coordinated descending turn to arrive
at the 90 degree position with 300 feet and 70
knots. Continue to decelerate and turn to arrive
on the wind line with approximately 1000 feet of
straightaway and 300 feet above the ground or 100
feet of altitude above the highest obstacle. Air¬
speed should be smoothly reduced to 45 KIAS as
the approach angle is reached. Reduce collective
and adjust cyclic to commence a descent on t the
desired approach angle. Keep the point of intended
landing in sight through the windshield. The
airspeed is controlled by nose attitude and the rate
of descent is controlled by the collective. Power
requirements are governed by the gross weight,
wind velocity, density altitude, and approach
angle. Commence gas producer acceleration and
slow the rate of descent with collective,
simultaneously reducing airspeed with cyclic so as
to arrive over the point of intended landing with 3
to 5 feet of altitude and zero airspeed. This should
be effected with little or no flare. The landing from
a hover is standard.

During steep approaches at less than 40
knots, avoid descent rates exceeding 800
FPM. See power settling paragraph in
Section 4.

High Speed Approach and Landing.

The high speed approach is employed to accelerate
the transition from flight to landing. Airspeed is
maintained in excess of 100 KIAS to an altitude of
100 feet, at which point the quick stop technique is
employed to transition to a landing. Rotor rpm will
tend to overspeed during the approach and quick
stop. Adjust collective and throttle as necessary to
maintain rotor/engine rpm within limits. Return
of throttle to the governor range must be effected
early enough to permit the engine to accelerate, so
as to arrive at 45 KIAS, level attitude in order to
transition from a steep approach to a landing.

CAUTION

Rapid application of throttle at or near
flat pitch can result in a rotor or engine
overspeed and subsequent damage.

-H1AAB-1 Section III

Part 2

Maximum Gross Weight Landing (No Hover
Landing).

Maximum gross weight landings should be
practiced to simulate landing without hovering at
high gross weights and high density altitudes.
This type of landing may be employed where a
transition to a hover is not possible or a sliding
landing is not feasible. The helicopter is flown as
in a normal approach with the exception that a
straightaway of 1000 feet, 70 KIAS, and 125 feet of
altitude are desirable. At this point raise the nose
attitude to slow airspeed and adjust collective to
slow the rate of descent. As the airspeed decreases,
continue to adjust collective to maintain a slow,
controlled rate of descent. As translational lift is
lost, level the helicopter and assume the landing
attitude. Continue to increase collective to
maximum power available to prevent a hard
landing. Touchdown should be at less than 5 knots
ground speed. Once the helicopter is firmly on the
ground, smoothly lower the collective to the
bottom to complete the landing. No hover landings
should be made, whenever possible, when
operating in sandy or dusty areas to minimize
wear on engine and rotor blades.

Sliding Landing.

Sliding landings should be practiced to simulate
conditions where hovering in ground effect is not
possible. They also aid the pilot in assessing the
feasibility of an operation requiring maximum
helicopter performance. They have value in that
they acquaint the pilot with the characteristics of
skid-type landing gear on various landing
surfaces and they afford the opportunity to
evaluate possible landing sites in case of engine
failure. If an emergency autorotative approach is
necessary, a sliding landing has the advantage of
greater helicopter controllability during
touchdown. It affords a safer landing with heavy
gross weights as well.

To practice sliding landing, select a firm, smooth
surface of sufficient length and free of
obstructions. The helicopter is flown as in normal
approach until just prior to touchdown. Maintain
sufficient forward speed to retain translational lift
and smoothly and slowly lower the helicopter to
the ground with the collective. Maintain heading
with pedals. Do not land the helicopter in a crab.
Compensate for any crosswind with the “wing
down” method. Landing attitude should be skid
level to prevent any pitching of the helicopter at
touchdown. Do not lower collective abruptly
during slide. Once the helicopter is on the ground,
allow the helicopter to slide to a gradual stop.
When the helicopter has stopped, lower collective
to the bottom.

Change 1 3.15

Section III
Part 2

NAVAIR 01-H1AAB-1

NOTE

Sliding landings on soft surface such as
mud, loose sand, and plowed fields may
cause the skids to dig in. This could
result in an abrupt stopping of the
helicopter, possibly causing severe
structural damage or a nose-over crash.

AUTOROTATION PRACTICE.

Full autorotation landings shall not be attempted
as a practice measure except by pilots specifically
authorized by competent authority. Practice
autorotations with power recoveries are permitted;
however, recovery will be initiated with sufficient
altitude to permit full recovery at an altitude of
3 to 5 feet above the surface. From this point a
waveoff may be accomplished straight ahead as in
a normal takeoff.

Practice autorotations will always be made into the
wind and will be performed at approved landing
areas or airfields. Always plan an autorotation to
an area that will permit a safe landing in an actual
emergency, preferably a hard, flat smooth surface,
clear of approach and roll-out obstructions.
Practice autorotations should not be attempted in
conditions of critical center of gravity loadings.
Caution should be exercised when practicing auto¬
rotations under conditions of high gross weight
because angle of descent is steeper, and rotor rpm
has a tendency to build up and is harder to control.
For practice autorotations, the minimum entry
altitude should be 500 feet above the ground and
not less than 70 KIAS for straight in autorotations.
The minimum entry altitudes will be 750 feet AGL
for 90 degree autorotations and 1000 feet AGL
for 180 degree autorotations. To initiate the
maneuver, reduce collective to the full down
position, simultaneously rolling off the throttles
to the engine idle detent. Check Ng at flight idle
and maintain heading and/or balanced flight with
pedals. During advanced phases of training,
deviation from minimum altitude and speed will
be at pilot’s discretion but not less than 100 feet
and 70 KIAS.

If the helicopter is only slightly out of
balanced flight, the rate of descent will
be increased by about 500 feet per
minute. An acute unbalanced condition
can result in an extremely high rate of
descent.

Adjust collective as necessary to maintain rotor/
engine rpm within limits.

Basic autorotation descents are performed at a
constant 80 KIAS and in balanced flight. At
approximately 75 to 100 feet above the ground,
commence a smooth flare, sufficient to slow the
airspeed and rate of descent. The rate and degree
of flare necessary will vary with airspeed, gross
weight, height above the ground, wind conditions,
and desired ground speed for landing. Adjust the
collective as necessary to keep rpm within limits.
Roll throttles open enought to join ENG RPM (Nf)
needles with ROTOR RPM (Nr) needle at 100
percent and increase collective slightly. Check GAS
PRODs to ensure that gas producers are accelerating
normally. Approaching the apex of the flare,
smoothly roll both throttles to the full OPEN
position. When the ground speed has been slowed
to a safe sliding landing speed, smoothly and
positively lower the nose of the aircraft to achieve
a skids level attitude by 20 feet. At approximately
15 feet of altitude, smoothly increase collective
and throttle to stop the descent at 3 to 5 feet
above the ground while maintaining 100 percent
ROTOR RPM (Nr) ENG RPM (Nf). Practice auto¬
rotations may be terminated to a hover or with
forward ground speed below 15 knots.

An excessive nose high attitude in the
flare at too low an altitude will result in
dragging the tail skid. This can cause
serious structural damage to the tail
pylon, possible tail rotor failure, and
uncontrolled flight.

3-16 Change 1

NAVAIR 01-H1AAB-1

Section III
Part 2

At average gross weights, best glide speed is
approximately 110 KIAS and minimum rate of
descent speed is approximately 65 KIAS.
Skidding/slipping the helicopter or reducing air¬
speed will increase the rate of descent and prevent
over-shooting. However, it is important that the
helicopter be returned to balanced flight prior to
commencing the recovery (flare).

HOVERING AUTOROTATION.

From a normal hover (not more than 5 feet), roll
off the throttles to flight idle, taking care not to
raise or lower the collective inadvertently. Use
sufficient right directional control pedal and right
cyclic to maintain heading and ensure a vertical
descent. The helicopter will tend to maintain
altitude momentarily, then will commence to
settle. As it settles, apply up collective to cushion
the landing. After the helicopter is firmly on the
deck, lower the collective to the full down position
and smoothly roll the throttles on to full open.

CAUTION

Hovering autorotations should only be
practiced at a moderate gross weight. At
heavy weights, greater skill and training
are required to cushion the landing and
there is a greater possibility of structural
damage to the helicopter.

DUAL ENGINE FAILURE (SIMULATED).

Simulated dual engine failures may be performed
in daylight, night, and under hooded instrument
conditions. They shall also be practiced over
suitable terrain for the purpose of developing
sound pilot judgement in the selection of the best
available emergency landing site. Wind direction,
air start procedures, and Mayday calls should also
be considered. Deviations from straight-in auto¬
rotations and varying entry airspeeds should be
practiced to ensure full utilization of the
helicopter’s capabilities and additional pilot
training. Simulated dual engine failures shall be
terminated no lower than 200 feet AGL and not
less than 60 KIAS.

i: caution ::

During simulated dual engine failures
initiated above 120 knots, if an aft
cyclic input is not made, a SCAS nose-
down pitch correction combined with
rapid decrease of collective could cause
less than +0.5 G loading which may
result in excessive main rotor flapping
and possible mast bumping.

QUICK STOP.

The quick stop is a maneuver used to reduce
airspeed as rapidly as safely feasible. It is useful in
aborting takeoffs, avoiding other aircraft or
transitioning from flight to an immediate landing
attitude. Execute a normal takeoff and accelerate
into forward flight. Establish stable 100 KIAS
flight at a constant altitude of 100 feet. For the
quick stop reduce collective and apply coordinated
aft cyclic to slow airspeed while maintaining
constant attitude (do not flare so abruptly that the
helicopter balloons). Adjust collective and throttle
as necessary to maintain rotor limits. When
airspeed has slowed to 45 KIAS, level the
helicopter and smoothly transition from a steep
approach to landing.

CAUTION

Reducing collective rapidly and
applying aft cyclic can result in rotor
overspeed.

TWENTY AND THIRTY DEGREE DIVES.

Twenty and thirty degree dives are practiced to
simulate high level rocket attack and to acquaint
the pilot with ordnance delivery maneuvers. These
dives should be initiated at or above 2000’ AGL to
have ample time and altitude for a smooth pulloff.
Set 40% torque and raise the nose 20 degrees above
the horizon. Slow to approximately 60 knots and
smoothly roll the aircraft toward the target line while
maintaining the nose on the horizon. Fifteen
degrees prior to intercepting the target line, begin
to roll wings level and allow the nose of the aircraft
to fall below the horizon. Stabilize on the target
line at the desired dive angle while maintaining
balanced flight and 40% torque. To avoid weapons
fragmentation, recover prior to 1000’ AGL by

Change 1

3-16A

Section III
Part 2

NAVAIR 01-H1AAB-1

raising the nose and then rolling the aircraft away
from the gun target line while simultaneously
increasing collective.

PRACTICE HIGH SPEED LOW LEVEL AUTO¬
ROTATIONS.

The practice high speed low level autorotation is a
maneuver used to simulate dual engine failure at
low level. To perform the maneuver, establish 100-
140 KIAS at a constant altitude not less than 100
feet AGL or 50 feet above the highest approach
obstacle. To enter the autorotation, reduce the
throttle to flight idle and apply coordinated aft
cyclic with collective reduction to slow the aircraft
and to maintain Nf/Nr within limits. At 80 KIAS
and 75 feet AGL, complete a normal autorotational
approach.

During practice autorotations initiated
above 120 knots, if an aft cyclic input
is not made, a SCAS nose-down pitch
correction combined with rapid decrease
of collective could cause less than
+0.5 G loading which may result in
excessive main rotor flapping and possible
mast bumping.

SINGLE ENGINE FAILURE (SIMULATED).

This maneuver will be performed in the training
environment to simulate a single engine emergency.
Simulated single engine failures shall be practiced
only when single engine flight, landing, or recovery
by autorotation is possible in the event of dual
engine power loss. Fly the landing pattern to arrive
at 500’ AGL and 80 knots at the abeam position.
Commence a coordinated descending turn to
arrive at the 90 degree position at 300’ AGL and
70 knots. Continue to decelerate and turn to arrive
on the wind line with a shallow glide slope. Slow
the aircraft while continuing descent to arrive at
the landing site. Procedural steps should conform
to single engine emergency procedures as stated in
Section V with simulation of appropriate steps
(2, 6, 7, 8, 9, 11). Single engine waveoffs may be
initiated as judgement dictates but should not be
attempted below 75’ AGL or 45 knots.

j: caution |j

It is important to closely monitor the
ITT and Ng of the good engine to pre¬
clude engine damage.

NOTE

Intentional single engine takeoffs are

prohibited.

MANUAL FUEL FLIGHT.

Manual fuel flight should be practiced to simulate
governor or automatic fuel control malfunctions
and to acquaint the pilot with the flight character¬
istics of the helicopter when operated in the manual
fuel mode. The maneuver will be initiated by
rolling both throttles to flight idle while on the
deck. The engine with the simulated malfunction
should then be switched to the manual mode.
Proper fuel control switchover should be verified
by a corresponding advisory panel indication and
slight Ng fluctuation. A “pop” which is caused
by a slight compressor stall may be heard when
switching into manual fuel. While maintaining the
throttle for the engine in manual fuel at flight idle,
increase the other throttle to full open. Increase
the throttle in manual fuel slowly to allow engage¬
ment of Nf and adjustment of engine torque. The
recommended technique for this maneuver is to
maintain the manual fuel torque needle at the
bottom of the transmission torque triangle. To
ensure that engine limits are not exceeded, make
all power changes slowly and smoothly during
manual fuel flight. To complete the maneuver, land
the aircraft and roll both throttles to flight idle.
Switch the engine, which is in manual fuel, back to
auto and verify the proper advisory panel indication
and Ng fluctuation. Slowly increase both throttles
to full open.

: caution :

• The Nf and AFCU overspeed governors
are not functional for the engine operated
in the manual fuel mode.

• In the event of an intermittent or
sustained electrical failure, the fuel
control will automatically revert to the
automatic mode. Should electrical
failure occur, place all governor control
switches in the automatic position
immediately.

3-16B

Change 1

NAVAIR 01-H1AAB-1

Section III
Part 2

TAIL ROTOR MALFUNCTION (SIMULATED).

The simulated tail rotor malfunction maneuver is
practiced to simulate landing with a fixed pitch
on the tail rotor. Once established at pattern
altitude, the pilot not actually flying the maneuver
will hold the pedals in a fixed position (the pilot
in control does not remove his feet from the
pedals). The aircraft will react to control inputs in
the following manner:

Collective increase
Collective decrease
Throttle increase —
Throttle decrease
Airspeed increase
Airspeed decrease

— yaw right

— yaw left
yaw right
yaw left
yaw left

— yaw right

The pilot may adjust collective, throttle, and
airspeed to become familiar with the aircraft
reaction to each adjustment and to determine how
I these responses may be used on final approach to
effect a landing. When possible, turns should be
| made in the direction of the yaw. The approach
should be flown with a slightly wider abeam
position, on a shallow glide slope, and with no
rapid power applications. Minimum control move¬
ments should be made on final. On short final,
adjust the collective and airspeed as necessary to
align the aircraft with the left side of the runway.

1 Continue a low approach to simulate a sliding
landing at not less than 3 feet AGL. A small
[ throttle reduction may be used to yaw the nose to
the left for final heading correction prior to a
simulated touchdown. At no time will throttles
be reduced below activation of the rpm audio
warning signal (in no case less than 92%). Waveoff
| should be executed in balanced flight at 100% Nr.

DUAL HYDRAULIC FAILURE (SIMULATED).

Dual hydraulic system failure (simulated) is a
maneuver practiced to enable the pilot to land the
helicopter in the event of a dual hydraulics system
failure. All approaches will be made to a simulated
sliding landing (low approach) at no less than 3
feet AGL and not less than 20 knots airspeed.
On the downwind leg at 500 feet and 80 knots, set
30% torque, turn off the No. 1 hydraulics system
and disengage the SCAS. Reset the Master Caution
light and check the No. 1 hydraulics gage for a
zero psi reading. 30% torque (with throttles full
open) is the minimum torque allowable by
I collective application, however, torque may be

increased up to 50% during the approach if
necessary. Extend the downwind to establish a
straight in approach of at least 2 NM. The abeam
position should be wider to allow for a shallow
turn. Throttles may be reduced to assist in
establishing a descent, but Nr must be maintained
at or above the limit for the rpm audio warning
signal (in no case less than 92%). Maneuver the
aircraft as necessary to establish a rate of descent.
At light gross weight configurations, the minimum
power setting (30% torque) may not result in a rate
of descent unless the airspeed is reduced to below
35 knots. Airspeed should not be reduced below
20 knots. Ensure that both throttles are full open
and the aircraft is aligned with runway heading
prior to commencing the simulated slide on landing.
The maneuver is terminated at no less than 3 feet
AGL and on the waveoff full systems will be
restored prior to turning downwind.

WAVEOFF.

Power-On Approach.

1. Collective - SMOOTHLY INCREASE TO
TAKEOFF POWER.

2. Airspeed - INCREASE TO CLIMB AIR¬
SPEED.

3. Cyclic — ESTABLISH A CLIMB.

Autorotative Approach.

1. Throttles - INCREASE TO FULL OPEN.
(Coordinate with collective to prevent
overspeed.)

2. Collective - SMOOTHLY INCREASE TO
TAKEOFF POWER.

3. Airspeed — INCREASE TO CLIMB AIR¬
SPEED.

4. Cyclic - ESTABLISH A CLIMB.

Change 1

3-16C

Section III
Part 2

SHUTDOWN.

NAVAIR 01-H1AAB-1

Pilots shall ascertain, prior to shutdown,
that the area is clear and that personnel
around the helicopter are outside the tip
path of the main rotor. Dming any
operations, the pilot is responsible for
keeping personnel around the helicopter
to a minimum number for safe operations.

1. Collective — DOWN.

2. Controls — CENTERED.

3. Canopy jettison pins — IN.

4. Throttles — ENGINE IDLE (61 ± 1%) FOR 1
MINUTE.

5. FORCE TRIM - ON.

6. ROTOR RPM AUDIO - OFF.

7. RADAR ALTITUDE altimeter — OFF.

8. KY-28 - OFF.

9. Radios and navigation equipment — OFF.

10. Countermeasures equipment — OFF.

11. ECU/VENT-OFF.

12. IDLE STOP — AFTER ONE MINUTE REL
- BOTH ENGINES.

Momentary actuation of the IDLE STOP REL will
result in the solenoid remaining retracted for a
period of five seconds. Thus the throttle may be
closed to the off position anytime within five
seconds after the switch is momentarily actuated.

13. Throttles — CLOSE.

14. SCAS POWER — OFF.

15. INVERTERS - OFF.

16. NO. 1 GEN, NO. 2 GEN - OFF.

17. Fuel — OFF (at zero percent).

CAUTION

•:

• Without the use of rotor brake on
shutdown, winds'of approximately 35
knots or above may cause the rotor to
windmill indefinitely (e.g., 20% ROTOR
RPM (Nr).

• If severe main rotor flapping or mast
bumping occurs due to high/gusty winds,
apply cyclic into wind, as required to
prevent or eliminate mast bumping.

18. Rotor brake — ENGAGE BETWEEN 60-25%
Nr.

NOTE

If rotor brake chatter or loud ticking
noise occurs, notify maintenance.

19. Lights — OFF.

20. BATTERY - OFF.

21. Collective strap — SECURE.

POSTFLIGHT EXTERNAL INSPECTION.

A postflight inpection should be made by the pilot
upon leaving the helicopter after completing the
assigned mission. This inspection is a general
visual inspection of the landing skids, fuselage,
tail rotor and drive systems, tail assembly and
engine compartment. In addition to the
established requirements for reporting any
systems defects, the pilot will also make entries on
the yellow sheet to indicate when any normal
operating limits contained in this manual have
been exceeded. When an emergency fuel is used,
report the type fuel and length of operation.

NOTE

Any contact with salt water spray shall
be noted on the VIDS/MAF.

3-16D

Change 1

NAVAIR 01-H1AAB-1

Section III
Part 2 — Part 3

NIGHT FLYING.

The procedures for night flying will be essentially
the same as those for days; however, visual
reference and depth perception are reduced.

Restrictions on Night Flying.

Helicopters shall not be flown at night if any of the
following equipment is not in operating condition.

»> ! ,

CAUTION

! ► ! ►

1 Ih w vhhwuwhw

1. Pilot compartment instrument and console
lights.

2. All exterior lights.

3. UHF radio.

4. Pilot gyro horizon.

5. Radar altimeter.

Helicopters shall not be flown beyond the
immediate vicinity of the field unless under
positive control of the tower if any of the
following equipment is not in operating condition:

1. Radio direction finding and navigation
equipment.

When landing in grass area, turn
searchlight OFF after landing to
prevent fire hazard.

2. Compass gyrosyn.

PART 3 — SHIP BASED PROCEDURES

COMMAND RESPONSIBILITY.

Shipboard environment, procedures, and
operations must be as normal as those used
ashore. The squadron is no longer an independent
command when embarked aboard ship but has
become an integrated part of an operating system.
Marine squadrons embarked tor amphibious
operations are component parts of the landing
force under the command of the Landing Force
Commander. The Amphibious Task Force
Commander exercises his command authority of
these units through the Landing Force

Commander. All squadrons embarked become a
part of the overall ships function for coordination,
control, and support. The commanding officer of
the squadron is responsible at all times for the
combat readiness of his organization. Command
relations and general procedures are contained in
NWP 42 and NWP 22-3.

FIELD CARRIER LANDING PRACTICE.

Field carrier landing practice (FCLP) is required of
all pilots within 30 days prior to carrier
qualification to ensure maximum crew

Change 1

3-17

Section III
Part 3

NAVAIR 01-H1AAB-1

proficiency. The number of periods will depend on
the experience and ability of the individual pilot
however, a minimum of two FCLP periods are
required (one day and one night period) FCLP’s
will be conducted to simulate shipboard operations
as closely as possible.

BRIEFING PRIOR TO FCLP.

1. Patterns, altitudes, and airspeeds.

2. Helicopter director signals.

Night FCLP.

When facilities permit, pilots should complete
FCLP s pnor to night carrier qualification to
familiarize themselves with night shipboard
landing procedures.

CARRIER QUALIFICATION.

The term carrier qualification referred to herein
encompasses all shipboard landing operations.
Initial day/night carrier qualification should be
made under ideal weather conditions including a
visible horizon.

Carrier Qualification And Requalification
Requirements.

Nothing in this* manual precludes the
commanding officer from exercising his own
judgment concerning the ability of a pilot to
perform a mission involving recovery on board or
when operational necessity dictates.

REQUIREMENTS. QU '^IFICATION

1. Day initial qualification: No less than 5
landings and takeoffs.

2. Night initial qualification: Day qualified and
not less than 5 night landings and takeoffs.
At least 2 day landings must be made on the
day of night qualification.

REQUALIFICATION REQUIREMENTS.

1. Day: Not less than 2 landings.

2. Night: Not less than 3 landings and at least 2
day carrier landings must be made on the day
of night qualification.

3. Currency: Requalify every 12 months.

4. If pilot has not met the requirements for
requalification in a 12 month period, subject
pilot is no longer current and must meet
initial qualification requirements.

LANDING AND RECOVERY PROCEDURES.

Shipboard qualifications are conducted using the
same procedures contained in the launch and
recovery operations in this section

Flight Scheduling.

Refer to NWP 42.

Briefing.

All pilots will receive a thorough briefing by the
ship’s Air Department Officer or his
representative on the ship’s air operations and
procedures. Flight briefings will be conducted by

o he w m mf , 0peration Department prior to each
Right. This detailed briefing will include the
information set forth in this section, and shall
include the following:

1. LSE.

2. Wind direction and velocity for flight
operations.

3. Use of helicopter lights (if night operation).

4. Traffic patterns and altitudes about ship.

5. CCA recovery and/or scheduled recovery
time.

6. Special safety precautions during shipboard
operation.

7. Ship s point of intended movement and
nearest land.

8. Aircraft deck spotting.

9. Ship’s navigational aids.

3-18 Change 1

NAVAIR 01-H1AAB-1

Section III
Part 3

10. Weather forecast and weather over nearest
land.

11. Ship’s position in the Force.

Hanger and Flight Deck Procedures.

S Deck procedures are found in CV NATOPS manual,
NWP 42 and NWP 50-2.

OPERATION OF EQUIPMENT.

Only qualified personnel shall operate towing
equipment. Towing couplings shall be inspected
prior to towing. Only approved tow bars will be
used. Ground handling wheels shall be installed in
eye bolts provided on each landing gear skid tube,
located forward of aft cross tube and forward of the
forward cross tube. Reference maintenance
manual for proper ground handling gear
installation and operation.

Care shall be taken to ensure that the
ground handling pins are properly
installed into eyebolts on the skid tube.

Ground Handling Gear Types.

Two types of ground handling gear can be used for
moving the helicopter, forward mounted and aft
mounted.

AFT GROUND HANDLING GEAR.

At gross weights of 13,560 pounds or less,
the aft ground handling gear may be used for
moving the helicopter. While in movement each
wheel assembly should be manned by a qualified
aircraft handler. A qualified aircraft handler shall
be positioned on the tail skid to take the weight off
the front of the skid tube and to provide steerage.
Two aircraft handlers may be utilized on the tail
skid when wind/weight conditions warrant. The
helicopter may be towed or pushed by hand if a
sufficient number of aircraft handlers are
available. Care should be exercised when lowering
the helicopter onto the deck. The helicopter should
be lowered slowly, and after assuring all personnel
are well clear of the helicopter.

FORWARD MOUNTED GROUND HANDLING
GEAR.

The forward ground handling gear should be used
when helicopter is at a high gross weight and/or
forward eg.

PROPER OPERATION WHEN FORWARD
MOUNTED GROUND HANDLING GEAR IS
USED.

Install all ground handling gears in eyebolts on
skid tube. Extend aft ground handling wheel on
one side only. Extend forward ground handling
gear on the same side. Extend remaining aft
ground handling gear. Extend remaining forward
mounted ground handling gear. For lowering
reverse the procedure.

Do not raise or lower forward mounted
ground handling gear unless the aft
ground handling gear is raised.

During actual movement of the helicopter each ■
hand brake shall be manned by a qualified aircraft I
handler. Hand brakes on ground handling wheels P
shall be applied immediately upon whistle or hand
signal. ■

FLIGHT DECK OPERATIONS.

1. Flight deck handling procedures and aircraft
handling signals are contained in NWP 42
and NWP 50-2.

2. Personnel not required for plane handling
will remain clear of the flight deck during
launch and recovery of helicopters.

3. Starting engine and rotor shall be done only
upon direction of personnel from the ship’s
Air Department.

4. Air taxiing and movement of helicopters shall
be under the positive control of LSEs.

3-19

Section III
Part 3

NAVAIR 01 -HIAAB-1

Manning Helicopters.

Upon receipt of the word to MAN AIRCRAFT,
flight crews will expedite movement to the
helicopter and complete the preflight inspection
and man aircraft.

Starting Engines and Rotor.

Preparations for starting the engine and rotor
shall be completed by the helicopter crew
immediately after they enter the helicopter.

Auxiliary power should be plugged in; the prestart
check list completed; inverter, UHF radio, TACAN
radio, FM radio and automatic direction finder on
for warmup, and a visual check of the surrounding
deck area should be made.

CAUTION

«; < *

Monitor voltage supplied by auxiliary
power at 26 to 29 vdc.

Mandatory requirements for starting engine and
rotor consists of the following items:

1. Main rotor and tail rotor blade tiedowns shall
be removed.

2. Offset main rotor blade to prevent tailboom
strike.

3. Deck tiedown secure.

4. Flight deck area clear of unnecessary
personnel.

5. Rotor engage/disengage wind limits (see
figure 3-3).

6. Fire guard on station.

0 Engines shall be started only on signal from a
LSE and under positive control of PRI-FLY.
Start procedure is normal. It may be necessary for
the pilot to adjust tip path plane. Cockpit checks
are accomplished in the normal sequence but
should be made as expeditiously as possible
consistent with safety.

Flight control checks, if necessary, shall
be performed with an absolute minimum
of flight control displacement. When the
helicopter is operating at 100 percent
FNG RPM (NO, a moderate amount of
right directional control pedals shall be
applied when the collective pitch control
is full down to prevent the helicopter
from skidding on the flight deck.

Tiedowns shall be removed when the pilot
signifies that he is ready for launch and the m
LSE has received permission to launch from B
PRI-FLY. The pilot will ensure complete removal
of tiedown chains prior to takeoff. In case of
downed helicopter, tiedown chains shall be left on,
and disposition of the helicopter will be
determined immediately after the launch. All
flight deck operations, including starting engine
and rotors, removing tiedown chains, etc., are
executed on signals relayed from PRI-FLY. The —
pilot should keep the LSE in sight and be prepared Q
to receive signals at any time.

Launch and Recovery Operations.

All commands are given by Primary Flight m
Control. LSEs relay all signals given by PRI-FLY g
when aircraft is in close proximity to flight deck.

RELATIVE WIND FOR LAUNCH AND
RECOVERY.

1. For launch and recovery wind limits, see
figure 3-4. In an emergency, the helicopter B
may be launched in 60 knot relative winds. Ml

2. Operations in the island wash area should be
held to a minimum.

LAUNCH PROCEDURES.

1. Helicopter shall not takeoff until cleared by

PRI-FLY and a signal has been received from *i
the LSE. |

2. Helicopters shall take maximum advantage
of available deck while gaining transitional
lift.

3-20

NAVAIR 01-H1AAB-1

AH-1T (Tow and Non-Tow) ROTOR ENGAGEMENT/DISENGAGEMENT WIND LIMITATIONS

HELICOPTER ALIGNED WITH SHIP'S CENTERLINE

SHIP HEAD
/ 000 \

FUNCTIONAL ROTOR BRAKE

LPH SPOTS 1-5
LH A SPOTS 1-7

Figure 3-3. Wind Limitations

Section III
Part 3

3-21

Section III
Part 3

NAVAIR 01-H1AAB-1

AH-1T (Tow and Non-Tow)

NO ROTOR BRAKE ENGAGE/DISENGAGE WIND ENVELOPE

HELICOPTER ALIGNED WITH SHIPS CENTERLINE

SHIP HEAD

LHA SPOTS 2-5
LPH SPOTS 2,3,4
SHIP ROLL 0-5°

Figure 3-4. Wind Envelope (Sheet 1 of 4)

3-22

NAVAIR 01-H1AAB-1

AH-1T (Tow and Non-Tow)

DAY LAUNCH/RECOVERY WIND ENVELOPE

HELICOPTER ALIGNED WITH SHIPS CENTERLINE

I I

SHIP ROLL 0-5°

LHA SPOTS 1-6
LPH SPOTS 1-4

FLIGHT OPERATIONS IN THIS AREA MAY REQUIRE LARGE,
RAPID YAW AND ROLL CONTROL INPUTS. APPROACHES,
LANDING AND TAKEOFFS SHOULD BE SLOW AND PRECISE.

FLIGHT OPERATIONS IN THIS AREA ARE CHARACTERIZED
BY LARGE POWER CHANGES AT THE DECK EDGE. THE PILOT
SHOULD ENSURE THAT THERE IS SUFFICIENT POWER
AVAILABLE FOR A NO WIND OGE HOVER IN THE EXISTING
AMBIENT CONDITIONS PRIOR TO A TAKEOFF OR LANDING.

Figure 3-4. Wind Envelope (Sheet 2 of 4)

Section III
Part 3

3-23

Section III
Part 3

NAVAIR 01-H1AAB-1

AH-1T (Tow and Non-Tow)

NIGHT LAUNCH/RECOVERY WIND ENVELOPE

HELICOPTER ALIGNED WITH SHIPS CENTERLINE

SHIP HEAD

I I

SHIP ROLL 0-5°

LHA SPOTS 4-7
LPH SPOTS 3,4,5

GSI ON OR OFF

FLIGHT OPERATIONS IN THIS AREA MAY REQUIRE LARGE,
RAPID YAW AND ROLL CONTROL INPUTS. APPROACHES,
LANDING AND TAKEOFFS SHOULD BE SLOW AND PRECISE.

Figure 3-4. Wind Envelope (Sheet 3 of 4)

3-24

iMAVAIR 01-H1AAB-1

Section III
Part 3

AH-1T (Tow and Non-Tow)

Days SCAS OFF Recovery Wind Envelope

HELICOPTER ALIGNED WITH SHIPS CENTERLINE

SHIP HEAD

SHIP ROLL 0-5°
LHA SPOTS 4-7
LPH SPOTS 3, 4, 5

FLIGHT OPERATIONS IN THIS AREA MAY REQUIRE LARGE,
RAPID YAW AND ROLL CONTROL INPUTS. APPROACHES,
LANDING AND TAKEOFFS SHOULD BE SLOW AND PRECISE.

Figure 3-4. Wind Envelope (Sheet 4 of 4)

Section III
Part 3

NAVAIR 01 -HI AAB-1

j: :

CAUTION

: : /

Moderate engine/rotor rpm droop and
slight settling of the helicopter may be
experienced immediately after liftoff
while clearing the deck. Transient droop
can be reduced by raising collective
slowly and smoothly.

3. Helicopters taking off will avoid crossing the
bow of the ship.

4. Rendezvous will be in accordance with
Section III, Part IV.

RECOVERY PROCEDURES.

STANDARD SIGNALS. Any of the following
standard signals may be given by flag hoist,
blinker and/or radio:

1. SIGNAL DELTA. The flight leader will orbit
his flight in the designated pattern.

2. SIGNAL CHARLIE. Commence landing.

DELTA PATTERN (HOLDING PATTERN). The
Delta pattern for helicopters is as designated in the
NWP 42 series or by the individual ship. More
than one Delta pattern may be designated.
This pattern may be assigned to any helicopter or
flights of helicopters during launch or recovery
operations. When helicopters are orbiting in a
Delta pattern, they will be prepared to break on
order from Primary Flight Control to join the
Charlie pattern.

LANDING PATTERN ENTRY/BREAKUP PRO¬
CEDURES. Unless cleared by PRI-FLY or direct
entry, helicopters shall approach the ship on a
heading that will parallel the ship’s base recovery
course close aboard the starboard side. The flight
leader starts his upward turn 400 yards ahead of
the bow at 300 feet altitude. Each succeeding
helicopter breaks to maintain a minimum but safe
interval.

CHARLIE PATTERN. This pattern is a race track
landing pattern oriented on the port side of the
ship and extending upwind a sufficient distance to
allow a normal landing interval between
helicopters. A designated altitude of 300 feet and
80 KIAS are maintained until starting the
approach to landing. On the downwind leg during
daylight operations, airspeed is reduced to arrive
at the 180-degree position at 70 KIAS, 300 feet of
altitude and about 400 yards abeam of the carrier.

I luring night operations, the 180-degree position is
70 KIAS, 300 feet of altitude and about 600 yards
abeam of the ship. 1 )epending on the relative wind,
the turn into the final approach is normally
started at a position even with the bow or just
ahead of the intended landing spot.

FINAL LANDING PROCEDURES. The
approach turn from the 180 degree position is
begun at 70 KIAS, adjusting speed as necessary to
maintain a rate of closure commensurate with the
relative speed of the ship. The final approach
should be relatively flat to eliminate the necessity
for exaggerated power changes or excessive flares
near the deck. Final movement of the helicopter
onto the deck position is normally accomplished
by forward and starboard movement of the
helicopter at air-taxi speed from the deck edge
about 8 to 10 feet above the deck. The pilot is
advised by signals from the LSE. Tiedowns shall be
attached prior to shutdown. Rotor RPM shall not
be decreased or engines shut down until signalled
by the LSE.

NOTE

A wave-off or hold signal from the LSE
is mandatory.

STARBOARD APPROACH. The starboard
approach may be authorized to facilitate ordnance
evolutions (including approaches with hung ord¬
nance) aboard LPH and LHA class ships. When
directed by PRI-FLY, enter the normal downwind
at 80 knots, 300 feet altitude and approximately
500 yards abeam the ship. Continue downwind to
a 180 degree position past the ship’s stern turning
to parallel the base recovery course about 50 yards
on the starboard side. Reduce altitude and airspeed
as necessary to permit a flat slide across the deck
edge in a controlled air taxi. During the final
portion of the approach, slide on an approximately
45 degree angle to the intended landing spot. The
aircraft should be landed on the spot parallel to the
ship’s centerline.

NOTE

On starboard approaches with unsafe
guns, hung or unexpended ordnance, the
final slide should be on a 30 degree angle
to the intended landing spot with
about a 15 degree nose right skid. The
aircraft should be landed on the spot
pointing to the ship’s 1 o’clock position.

EMERGENCY PROCEDURES. Any helicopter
experiencing trouble in flight will immediately
notify the flight leader by radio or by visual
signals as the situation dictates.

3-26

NAVAIR 01 -HI AAB-1

Section III
Part 3

If the nature of the emergency warrants an
immediate return to the ship, a radio call will be
made to enable the ship to prepare for landing. In
any case, the following information will be
transmitted to the ship:

1. Side number of the helicopter.

2. Position.

3. Difficulty.

4. Intentions.

If the helicopter having the emergency does not
have a radio contact with the parent carrier, all
possible information is relayed visually to the
wingman who makes the necessary radio
transmission. If communications are lost, the
helicopter signal to indicate an emergency is as
follows: Turn NAVIGATION LTS to FLASH and
landing light to ON during the approach to the
carrier.

AIR CAPABLE SHIP OPERATIONS.

I Air capable ships include all ships other than
aviation ships (CVs) and amphibious aviation
ships (LPHs/LHAs), for example, LPD, LSD,
LCC, LKA, DD, CG, etc. Basic shipboard pro¬
cedures used on LPH and LHA class ships
normally apply to operating on air capable ships
as set forth in NWP 42. Pilots should be aware that
except for LPDs, air capable ships have no air
department and will have little experience
operating with Marine helicopters. LPDs have two
landing spots and all other air capable ships with
a landing capability have one. Specific ship
helicopter capabilities including obstruction and
specific restriction are included in “Air Capable
Ship’s Helicopter Facility Resume” (NAEC-ENG-
7576).

Launch Procedures.

The LSE will signal for launch to either port or
starboard depending on obstructions and relative
wind (usually with 30 degrees of the ship’s
heading). Following the LSEs signal to lift, the
pilot will lift into an 8 to 10-foot hover and depart
the ship using one of the following methods:

1. Slide perpendicular to the ship’s centerline
over the deck edge to a minimum of one
rotor diameter from the ship; then transi¬
tion to a normal climb into the wind.

2. When wind conditions are favorable and
there are no obstructions to flight; tran¬
sition from a hover, parallel the extended
lineup line, and climb straight away from
the ship (approximately 45 degrees from
the ship’s heading).

NOTE

Special care must be exercised to
remain clear of crane and deck edge
obstructions.

Recovery Procedures.

Helicopters will be recovered individually from the
designated delta pattern. Landing patterns are
normally flown at 300 feet and 80 knots. The final
approach should be relatively flat to avoid flares
in the immediate vicinity of the ship. Normal
recovery procedures are as follows:

1. Obtain landing clearance and wind condition
from the ship.

2. At the abeam position, commence a normal
approach. Fly the approach down the
approach line to arrive just short of the
deck edge about 10 feet above flight deck
level with a gradual transition to an air taxi
condition; continue over the deck edge over
the approach/lineup line; land with the skids
in the center of the circle and the axis of the
aircraft over the lineup line. Touchdown will
be smooth.

NOTE

Due to the potential turbulence caused
by the ship’s superstructure and/or flight
deck edge, and slow airspeed while
crossing over the deck edge, gross'
weights should be limited to HOGE.

Stabilized Glide Slope Indicator (SGSI).

SGSIs are designed for use on air capable ships to
provide a visual assist to helicopter approaches,
including operations at night and during conditions
of reduced visibility. These units are being
installed on some surface ships.

NOTE

SGSI pilot procedures are different from
those used with the standard Marine
glide angle indicator light (GAIL),
therefore reference must be made to
NWP 42 for proper SGSI procedures.

3-27

Section III

NAVAIR 01-H1AAB-1

AH-1 LIGHTS

Condition

Ship Red Deck Lighting

Ship White Deck Lighting

Ready for external power

As required

Ready to start engines

Nav lights FLASH DIM

Ready for takeoff

Nav lights STEADY DIM

Nav lights STEADY BRT
! Anticollision light ON

After takeoff

Nav lights STEADY BRT
Anticollision light ON

Established downwind
prior to 180° position

Nav lights STEADY DIM
Anticollision light OFF

After final landing or
holding on flight deck

Nav lights STEADY DIM
Anticollision light OFF

Anticollision light OFF

Figure 3-5. Lighting Procedures ntsa so

NIGHT OPERATIONS.

Preflight Procedures.

The pilot’s red-lensed flashlight will be used in
making the external inspections. In addition to the
normal cockpit inspections, ensure that all light
switches are positioned properly. Lighting at night
becomes a critical area, and the general rule of not
showing white lights on the flight deck should be
rigidly observed.

Helicopter Lighting.

□ The helicopter lighting procedures (figure 3-5) shall

□ be used for all night shipboard operations.

Taxi and Operations.

The first rule the pilot .should remember con¬
cerning night shipboard operations is that the

tempo of operations, both in volume and speed is
considerably reduced from day operations. Slow
and careful handling of helicopters by both
helicopter directors and pilots is mandatory.
When a pilot has doubts about an LSE’s signal, he
should hold his position and request confirmation
from PRI-FLY.

Postflight Procedures.

Postflight procedures and the postflight inspec¬
tion are performed in the same manner with the
same caution concerning night visibility as is
required for preflight operations.

DEBRIEFING.

When based aboard ship, debriefing can be equally
as beneficial as that required when based ashore.
For detailed debriefing, refer to this section.
Debrief those portions applicable to the flight.

3-28

NAVAIR 01-H1AAB-1

Section III
Part 4

PART 4 — SPECIAL PROCEDURES

FULL AUTOROTATION LANDING.

A full autorotation landing is performed in the
same manner as the practice autorotation with
power recovery with the exception that the
throttles remain in the engine idle detent when the
flare is commenced. At about 10 to 12 feet of actual
altitude, smoothly raise the collective pitch control
to slow the rate of descent, apply sufficient forward
cyclic control to level the helicopter, and maintain
heading with directional control pedals. Do not
land in a skid. At about 2 feet altitude increase the
rate of collective pitch control movement so as to
effect a gentle touchdown. When the helicopter is
on the ground, stop collective pitch control
movement and allow the helicopter to slide to a
gradual stop, maintaining heading. The
touchdown will be made in a near level attitude, to
prevent adverse pitching of the helicopter.

CAUTION

Zero airspeed autorotative landings
should be avoided except in actual
emergencies when the available landing
surface is unsuitable for a sliding
landing.

FORMATION AND TACTICS.

Introduction.

It is essential that the basic fundamentals of
formation flying be practiced in preparation for
combat readiness. The procedures and positions
contained herein are intended to provide a
foundation for formation flying which will meet
most mission requirements, both combat and non¬
combat.

The signal for a change in a formation may be
accomplished by the use of the radio on a squadron
common frequency or appropriate hand signals as
contained in NWP 50-2.

In any case, no changes in the formation will take
place until all aircraft in the formation understand
and acknowledge the signal.

Formations.

ELEMENTS OF A FORMATION.

The number of aircraft required to accomplish a
mission varies. A section will consist of two
aircraft, and a division will consist of three or four
aircraft (two sections). Two or more divisions
constitute a flight. The disposition of members
within a formation is at the discretion of the
leader.

BASIC FORMATIONS.

The two basic types of formations are parade and
tactical. Parade is used primarily when there is a

Change 1

3-29/(3-30 blank)

NAVAIR 01 -HI AAB-1

Section III
Part 4

A - LEAD (FORMATION) AXIS
B - 45° BEARING EITHER
SIDE OF LEAD AXIS
C- HORIZONTAL SEPARATION
OF ONE ROTOR DIAMETER
10 FEET STEP-UP

2049L7-53A

Figure 3-6. Fingertip Parade

requirement of aircraft to fly a fixed bearing
position in close proximity to each other and
maximum maneuverability is not essential. It is
most frequently employed during arrival at or
departure from ships or airfields, or during flight
demonstrations. Power is varied to maintain
position. Maneuverability is a prime consideration
for formations engaged in combat tactics. The
leader must be able to use his formation as an
integral unit and still be free to turn, climb, and
dive the formation with few restrictions. The
tactical formations outlined herein afford this
flexibility. Radius of turn is varied rather than
power to maintain position.

Parade Formations.

TYPES.

The four basic types of parade formations are:
echelon, fingertip (figure 3-6), diamond (figure 3-7),
and column.

POSITIONS.

The parade position for echelon, fingertip and
diamond is on a 45 degree bearing either side of
lead axis with 10 feet of step-up, and one rotor
diameter diagonal clearance. This position
provides adequate longitudinal and lateral
clearance between aircraft. In fingertip and
diamond, the section leader will fly the same
position on the leader as the number two man. The
column position is on a 0 degree bearing with 10
feet of step-up and two (2) rotor diameters
longitudinal clearance.

PARADE TURNS.

ECHELON AND FINGERTIP. Wingmen will
rotate about the leader’s longitudinal axis during a
turn into them, and on their own longitudinal axis
on turns away from them.

3-31

Section III NAVAIR 01 -HI AAB-1

Part 4

Figure 3-7. Diamond Parade

DIAMOND AND COLUMN. Wingmen will
maintain a fixed position and roll about the
leader’s longitudinal axis on all turns.

CROSSOVERS.

Crossovers shall be accomplished by individual
wingmen, or sections when directed by the leader.
The leader shall ensure that all helicopters in his
formation are aware of the change in formation.
The following procedures will be followed:

1. When a wingman is required to crossover, he
will move to the corresponding position on

the opposite side maintaining longitudinal
blade tip clearance constant" The section
leader will slide out on bearing allowing room
for the number two aircraft when applicable.

2. When the section is required to crossover, it
shall be accomplished by the section moving
across to the appropriate position on the
opposite side. The section leader’s wingman
will not affect his crossover on the section
leader until the section leader is in his new
position.

3-32

NAVAIR 01 -HI AAB-1

Section III
Part 4

LEAD CHANGES.

All changes of the lead position i\i a formation shall
be acknowledged by the recipient in such a manner
as to preclude the possibility of misunderstanding
by any member of the formation. Preferably, a lead
change will be executed from level flight and in
such a manner as to allow the old leader time to
assume his new position before maneuvering flight
is commenced. The old leader shall maneuver to
establish step up and maintain one rotor diameter
separation while he moves back to his new position.

Tactical Formations.

TYPES.

The three basic types of tactical formations are:
cruise (figure 3-8), tail chase, and combat spread.

CRUISE.

The cruise position is on a 30 degree bearing off the
tail with 10 feet of stepup, and two (2) rotor
diameters diagonal clearance. This position will
provide adequate longitudinal and lateral
clearance between aircraft for maximum
maneuverability. Number three will fly a position
to allow room for number two between himself and
the leader. When the leader initiates a turn,
aircraft will maintain longitudinal clearance on
the aircraft directly ahead by sliding and utilizing
the radius of turn created by the leader. To
decrease distance, increase bank; to increase
distance, decrease bank. As soon as the leader rolls
level the normal cruise position will be resumed
with the No. 2 aircraft balancing the formation.

TAIL CHASE.

The tail chase formation is flown in a loose column
and all wingmen will utilize basic cruise-turn
principles to maintain longitudinal clearance
between aircraft. The formation is most commonly
used in conjunction with the basic cruise
formation when the leader is required to maneuver
extensively through climbs, dives, and turns for
the purpose of defensive evasive maneuvering,
approaches to confined lz’s, etc. When the need for
the tail chase ceases, the leader should return the
formation to normal cruise formation.

COMBAT SPREAD.

Combat spread is the basic formation used by
flights of armed helicopters in the conduct of
their mission. Variations of combat spread are used
as a basis for establishing escort formations.

Maneuvering in combat spread is accomplished
using the basic fluid four techniques (radius of turn
and altitude advantage are used to maintain/
regain position in the flight). Exceptions to the use
of fluid four are necessary when a flight is required
to execute a break. Refer to NWP 55-3-AH-l,
Vol. I.

RENDEZVOUS.

The two types of basic rendezvous are the running
rendezvous and the carrier type rendezvous. A
combination of the principles of these two is most
commonly employed to join aircraft after takeoff.

Running Rendezvous.

The leader will depart maintaining a prebriefed
airspeed and will allow wingmen to use an
airspeed differential and/or radius of turn that
will enable them to overtake the leader and join as
briefed.

Carrier Type Rendezvous.

Basically this is a join-up executed while the
division leader makes a 180 degree level turn,
using a 10 degree to 15 degree angle of bank and
100 KIAS. Joining helicopters will assume a
rendezvous bearing on the division leader using
the cut-off vector to affect the join-up. The final
phase of the rendezvous will be on a 45 degree
rendezvous bearing. Join-ups will be made to the
inside of the turn. After relative motion is stopped,
affect a crossover.

When practicing carrier type rendezvous, break¬
ups will be executed from an echelon. The leader
will break maintaining altitude and a 30 degree
bank throughout his turn. Each succeeding
wingman breaks at a prebriefed interval with 30
degree of bank, adjusting his bank to be in an
extended column position when the leader
completes his turn.

FORMATION TAKEOFFS AND
LANDINGS.

Formation takeoffs and landings are frequently
used during normal missions and should be
practiced.

Power available, size of zone, obstacles to flight,
wind direction and velocity, enemy fire, terrain
features, rotor turbulence, and other
considerations will determine the positions to be
assumed by members of a formation.

3-33

Section III
Part 4

NAVAIR 01 -HI AAB-1

204947 - 541 )

Figure 3-8. Four Plane Division Tactical Cruise

3-34

NAVAIR 01 -HI AAB-1

Responsibilities.

Section and division leaders must endeaver to fly
as smoothly and as steadily as possible,
maintaining constant altitudes, headings and
power settings. Section and division leaders are
responsible for maintaining positions within the

r Section III
Part 5

formation as instructed. The leader is responsible
for briefing, conduct, and discipline of the flight.
He normally handles radio transmission for the
flight, including takeoff and landing clearance.
All section leaders must be prepared to assume
lead of the division.

PART 5 — FUNCTIONAL CHECKFLIGHT PROCEDURES

INTRODUCTION.

Check Pilots.

Commanding officers will designate, in writing,
those pilots within their command who are
currently eligible to perform this duty.

Checkflights and Forms.

Checkflights will be performed when directed by,
and in accordance with, OPNAVINST 4790.2
series and the directions of NAVAIRSYSCOM
type commanders, or other appropriate authority.
Functional checkflight requirements and
applicable minimums are described below.
Functional checkflight checklists are promulgated
separately.

REQUIREMENTS.

Conditions Requiring Functional
Checkflights.

Checkflights are required under the following
conditions (after the necessary ground check and
prior to the release of the helicopter for operational
use):

A. At the completion of helicopter rework and
all phase 1) inspections (all checkflight
items required are prefixed A).

B. After the installation of an engine or engine
fuel control, or any components which
cannot be checked in ground operations
(minimum required are prefixed B).

C. When fixed or movable flights surfaces or
flight control system components have been
installed or reinstalled, adjusted, or rerigged
and improper adjustment or replacement of
such components could cause an unsafe
operating condition (minimum required are
prefixed C).

PROCEDURES.

Functional Checkflight.

The following items provide a detailed description
of the functional checks sequenced in the order in
which they should be performed. In order to
complete the required checks in the most efficient
and logical order, a flight profile has been
established for each checkflight condition and
identified by the letter corresponding to the
purpose for which the checkflight is being flown;

i.e., A through C above. The applicable letter
identifying the profile prefixes each check both in
the following text and in the functional
checkflight checklist.

FUNCTIONAL CHECKFLIGHT CHECKLIST.

Checkflight personnel shall familiarize
themselves with these requirements prior to the
flight. NATOPS procedures shall apply during the
entire checkflight unless specific deviation is
required by the functional check to record data or
ensure proper operation within the approved
aircraft envelope. A daily inspection is required
prior to the checkflight. Checkflight pilots shall be
briefed by Maintenance Control or Quality
Assurance personnel prior to flight.

Before Preflight.

The check pilot shall ensure the following have
been accomplished:

1. All discrepancies have been signed off by the
inspector; a qualified plane captain has
signed off the preflight and a responsible
authority has signed the helicopter off as safe
for flight.

2. The purpose of the flight portion of the check
card has been properly filled out.

3. There is no doubt what is required for a
complete and accurate check.

3-35

Section III
Part 5

NAVAIR 01-H1AAB-1

4. The plane captain has unbuttoned each
portion of the helicopter that is accessible to
preflight. Keep in mind the most important
aspect of a checkflight is the PREFLIGHT
and POSTFLIGHT.

Do not begin maintenance flight
readiness inspection until armament
systems are determined to be safe.

PROFILE.

ABC

Exterior Check.

1. Main rotor blades — Visually check condition and cleanliness.

2. TSU — Security, cleanliness.

3. Emergency canopy jettison — Secure, unobstructed.

4. Cover — Condition, all screws installed.

5. Stress panels — Secure, all screws installed.

6. Gun turret - Dzus fasteners secured, skin damage, gun locked in stow position no

excessive play in barrel. ’

7. Copilot/gunner windshield and window - Check for heat damage, scratches
evidence of leakage.

8. Rain removal — Unobstructed.

9. Static ports — Check for damage, obstruction (i.e., painted over).
10. Stress panels — Right side secure, all screws installed.

U. Ammunition bay — Trays installed properly with security pins in place. White
fatehing of C ba“ oo^ “ H “ da! " age ' C ° nditi °"’ P">P«

12. Pilot canopy door - Check for excessive scratches, damage or discoloration
proper tit, handles for damage, hinges and screws secured.

13. ADF and FM - Sensing and homing antenna for damage and proper installation.

3-36 Change 1

NAVAIR 01 -HIAAB-1

Section III
Part 5

14. Hydraulic compartment — Lines and fittings for security and leakage, all bypass
button indicators in, reservoirs full, ECU duct for condition, and door latches for
condition and operation. Gravity refueling cap — Secure.

15. Right wing — Surface condition, navigation and formation light for cracked glass
and security. Ensure safety levers are locked and safety pins in. Check slip marks
on 4 mounting bolts. Check tiedown ring is secure.

16. Compartment under wing — Check lines for leakage, looseness, cleanliness,
chaffing; fore/aft servo and synchronized elevator control rods for security and
leakage.

17. Landing gear — Check fairing for damage; fuselage for wrinkles (evidence of hard
landing); cross tubes for bends; skid shoes for proper installation, damage and
security; weather stripping for installation and security.

18. Transmission area — No foreign material, check oil leakage, check mounts for
lockwire, wear or damage, ensure control tubes are not rubbing lines or airframe,
dampeners, and lines for fraying.

19. Rotor brake disc — Condition, security.

20. Transmission oil jets — Installed, safetied.

21. Power cylinder mounts, hydraulic pumps, and tachometer generator — Secure.

22. Lift link — Check for cracks, bends, security of attaching points.

23. Tail rotor driveshaft — No grease leakage or scratches.

24. Main driveshaft — Proper alignment, not throwing grease, check drain line from
input quill. Check proper operation of free wheeling unit (rotate to port side).

25. Transmission oil filler cap and sight gage Security of cap and proper oil level.

26. Transmission chip detector plug and drain — Securty, leakage, and lockwire.

27. Accessory gearbox area — Check throttle linkage and all accessories for
condition. Check for fuel and oil leakage.

28. Engine intake area — Check for cleanliness, security, damage, and foreign matter.

29. Engine oil tank — Proper oil level, leakage, filler cap security, and chip detector
lockwired.

30. Transmission cowling — Check fasteners, hinges for cracks and proper fastening.
Check number 1 hydraulic filter buttons.

31. Engine compartment — Check for fuel and oil leakage; all lines and wiring for
tightness, chafing, leakage, and security. Check Nf governor linkage for security.
Check oil cooler blowers for security and foreign matter. Check engine mount.
Check hanger bearings.

3-37

Section III NAVAIR 01-HI AAB-1

Part 5

32. Transmission oil filters — Bypass indicator in.

33. Combining gearbox — Check for oil level, leakage, and security of chip plug.

;14. Exhaust pipe — Note any evidence of oil and check for cracks. Check for security of
thermal cover.

do. Combustion chamber area — Check for fuel leakage; condition of fire detector; no
obstruction to fire extinguisher.

36. Exhaust extension — Check for cracks, chafing, or looseness.

d7. Engine cowling — Check fasteners and hinges for cracks and proper fastening.

37a. Pressure refueling receiver cap — Secure.

38. Electrical compartment — Battery for security, cleanliness, evidence of corrosion,
leakage, vent lines clear, and lockwired. Tail rotor servo for leakage, all lines and
wires for tightness, tailboom mounting bolts for looseness (slip marks), all circuit
breakers in, and inverters for proper installation.

39. Tailboom (right side) — Skid condition, popped rivets, structural damage, all
access panel fasteners installed.

40. Driveshaft cover (right side) — Condition and security of skin and hinge.

41. Synchronized elevator (right side) — Skin condition and excessive play in spar.
Check trailing edge for separation.

42. 42 degree gearbox — Security, grease and oil leakage; filler cap and cover secure.
No dzus fasteners absent or cracks in cover. Check oil level.

43. Tail skid — No excessive play.

44. Aft navigation lights — Check for loose rivets, cracked glass, and condition.

45. Vertical fin (right side) — Skin condition, loose or popped rivets.

46. Tail rotor blades — Check for damage, freedom, free to flap, and tiedown removed.

47. Tail rotor hub and components — Check for excessive looseness of the yoke and
crosshead bearings, counterweights for correct positioning, pitch change links for
correct installation and wear, check appropriate components are lockwired.

48. 90 degree gearbox — Oil level, leakage, filler cap — secure.

49. Vertical fin (left side) — Skin condition, loose or popped rivets.

50. Vertical driveshaft cover — Hinges and fasteners for security.

51. 5th driveshaft — Scratches, dents, condition of 42 degree gearbox driveshaft fan.

52. 2nd, 3rd, and 4th driveshaft — Associated hanger bearings and clamps for grease,
security, nicks, dents, scratches.

3-38

NAVAIR 01-H1AAB-1

Section III
Part 5

53. 42 degree gearbox — Check for oil leakage and level.

54. Driveshaft cover (left side) — Condition, cracks, dzus fasteners installed and
secured.

55. Tailboom (left side) — Skin condition, popped rivets and structural damage, check
underside for breather screens attached and screws in place.

56. Synchronized elevator (left side) — Skin condition and excessive play in spar.
Check trailing edge for separation.

57. Electronic access panel — Check condition of doors and ensure electronic
equipment is secure with wires attached properly.

58. Exhaust pipe — Note any evidence of oil and check for cracks.

59. Exhaust extension — Check for cracks, chafing, and looseness.

60. External power receptacle — Security of unit, prongs, and door secured.

61. Fire extinguisher and radio access panel — Check condition of door, pressure in
fire bottles (550-700 lbs.), red indicators not discolored, check linkage and droop
cam, wires for security.

62. Engine compartment — Check for fuel and oil leakage, all lines and wiring for
tightness, chafing, and security, check engine mounts for axial play and cracks.
Check combining gearbox oil filter bypass indicator.

63. Engine cowling — Check fasteners and hinges for cracks and proper fastening.

64. Accessory gearbox — Check throttle linkage and all accessories for condition.
Check for fuel and oil leakage.

65. Engine oil tank — Proper oil level, leakage, filler cap secure.

66. Transmission — Lift link, main driveshaft, mounts, lockwire, rotor brake, mount
and bolts for cracks.

67. Transmission cowling — Check fasteners, hinges for cracks, proper fastening,
and pitot tube.

68. Number 1 hydraulic gage — Check level.

69. Pylon access doors — Check hinges and operation. Check number 1 hydraulic
reservoir and filter for leakage and condition.

70. Drive links — No excessive looseness.

71. Anti-drive link — No excessive looseness.

72. Mast boot — Check for security and damage.

73. Friction collet — In place, secure.

3-39

Section III
Part 5

NAVAIR 01 -HIAAB-1

74. Segmented clamp — Secure.

75. Scissors assembly — No excessive looseness.

Bolt attaching control tube to scissor
assembly must be installed opposite to
direction of rotation.

76. Lower bearing — Check for excessive looseness.

NOTE

Particularly check scissors area for
wear.

77. Swashplate — No vertical looseness or visual wear. Rotating swashplate — Check
for lockwire and condition.

78. Collective sleeve hub — Check for security and condition.

79. Static stops — Check for evidence of mast bumping and check attaching bolts for
shear offset.

80. Mast — Check for scratches, nicks, dents.

81. Upper bearings — No excessive looseness.

82. Pitch change rod and barrel — Scratches, dents, lockwire, jamnuts secure.

83. Main rotor hub — Check for damage and corrosion.

84. Grips — Check oil level and leakage.

85. Pitch horn — Check rod end attachment for looseness of bolts, nicks and
scratches.

86. Drag brace — Check condition and locknuts secure.

87. Trunnion and bearing — Check for looseness of bolts.

88. Blade attachment bolts — Condition, locks installed.

89. Mast nut — Secure with lock in place.

90. Main rotor blades — Check for bonding separation, cracks, cleanliness.

91. Upper cowling and anti-collision light — Check for damage, security, and
operation.

3-40

NAVAIR 01-H1AAB-1

Section III
Part 5

ABC

92. Left wing — Surface condition, position and formation light for cracked glass and
security. Ensure safety lever is locked and safety pin is in. Check slip marks on 4
mounting bolts. Check that tiedown ring is secure.

93. Compartment under wing — Check leakage, lateral servo, collective servo, lines
for chafing.

94. Landing gear — Check fairing for damage, fuselage for wrinkles (evidence of hard
landing), crosstubes for bends, skid shoes for proper installation, damage, and
security, weather stripping for installation and security.

95. Hydraulic compartment — Check ECU, access door and latches, rotor brake
cylinder for security. Check transmission chip detector lights.

96. Stress panels — Condition.

97. Ammunition bay — Trays installed properly with security pins in place. White
teflon slides and door cables for damage. Condition, operation, and proper
latching of bay door.

98. Pilot window — Check for scratches, discoloration.

99. Copilot/gunner canopy door — Check for scratches, discoloration, proper fit,
handle and hinges for damage.

100. Static ports — Check for damage or obstruction.

Pre-Entry Inspection.

As stated in NATOPS normal procedues; Section III, Part 2.

Interior Inspection (pilot station).

1. First aid kit installed, unopened.

2. Condition of det cord.

3. Check all gages for limit marks and proper installation.

4. Proceed with interior inspection as stated in NATOPS normal procedures, Section
III, Part 2.

Start.

1. After electrical power has been applied to helicopter, the following lights should be
illuminated on the master caution panel:

ENG 1 OIL PRESS, ENG 2 OIL PRESS

90° TEMP/PRESS, 42° TEMP/PRESS

ENG 1 PART SEP OFF, ENG 2 PART SEP OFF

NO. 1 DC GEN, NO. 2 DC GEN
XMSN OIL PRESS

3-41

Section III
Part 5

NAVAIR 01 -HI AAB-1

C BOX OIL PRESS

NO. 1 HYD PRESS, NO. 2 HYD PRESS
AC MAIN
XMSN OIL BYP

KW1) FUEL BOOST, AFT FUEL BOOST

EXT PWR DOOR OPEN (only if external power door is open)

ALT ENCODER (light disabled if AAU/32 is installed)

ABC
ABC

2. Engine 1 — Check engine idle (62 ± 2% Ng).

3. Engine 2 — Check engine idle (62 ± 2%Ng).

4. Flight controls — Refer to page 3-10

ABC

5. Force trim — Refer to page 3-10

6. Hydraulic check — Refer to page 3-10

7. Acceleration check — Throttle both engines to engine idle and check RPM switch
full decrease (engine idle 61 +1% GAS PROD (Ng)). Rapidly open No. 1 engine
throttle monitoring INLET TEMP and GAS PROD (Ng) speed. When GAS PROD
(Ng) passes 90%, rapidly close throttle to engine idle. Time for acceleration from
engine idle to 90% GAS PROD (Ng) should be 5 seconds maximum.

8. ENG RPM (Nf) Governor — Differential trim check — With throttles full open,
actuate ENG TRIM switch to full minus and note engines torque spread (#2
torque lower than #1). Then actuate ENG TRIM switch to full plus and note
engines torque spread (#2 higher than #1); the torque differential should be
approximately equal from full minus to full plus. Dual beep check — With
throttles full open, actuate ENG TRIM switch to match engine torques; then with
RPM switch,check dual governor range at 97% RPM (Nf) full decrease to 101.5
±0.5% ENG RPM (Nf) full increase. Running the governor through full range
should take 5-10 seconds. Engine trim — With throttles full open and ENG
RPM (Nf) set at 100%, pull in sufficient collective to obtain a light on the skids
condition. Check that the torque of the engines are within 3% of each other.
Check TRIM switch for a minimum of ± 3%.

a. No. 1 engine idle and No. 2 engine full throttle.

b. Record maximum RPM for No. 2 engine.

3-42

NAVAIR 01-H1AAB-1

Section Hi
Part 5

c. Repeat above for No. 1 engine.

d. Single engine maximum rpm should be between 98-100% ENG RPM (Nf).

AB 9. T5 BIAS - With one engine 700°C or above INLET TEMP, activate BIAS TEST

switch to selected engine and record the change in INLET TEMP. It should be the
data plate bias ± 5°C. If not, adjust accordingly. Repeat process for the
other engine.

CAUTION

I--1

A bias greater than that data plate bias
could result in an inadvertent overtemp.

AB 10. Ground power assurance check — (on deck) Perform the ground power assurance

check with the helicopter headed within 5 degrees of the prevailing wind.

a. Set 29.92 in altimeter and record altitude.

b. Determine and record torque required.

c. Check tower for current OAT in degrees C and winds. Check helicopter FAT
gage. Record both temperatures. Use the lower temperature or indicate
temperature used.

d. Determine and record max allowable INLET TEMP and GAS PROD (Ng).

e. Roll No. 1 throttle full open. Pull in torque determined in b above. Maintain 97%
ENG RPM (Nf).

f. Allow 5 minutes stabilization time and record:

(1) GAS PROD (Ng) to nearest tenth.

(2) INLET TEMP to nearest 5 degrees.

(3) ENG TORQUE.

AB 11. Particle separator, (Engine RPM) warning — With engines above engine idle

move PART SEP switch of No. 1 engine to the OFF position and check MASTER
CAUTION light illuminates (PART SEP OFF). Move switch to the ON position
and check light is out (there will be a slight delay before light goes out). Move
switch to AUTO position and check that light does not illuminate. When the
TRIP TACH or RPM WRN circuit breakers are out, the particle separator is
inoperative in the automatic mode. Repeat above for No. 2 engine. Reduce
throttle below engine idle, record GAS PROD (Ng) when respective PART SEP
OFF and MASTER CAUTION and RPM lights illuminate.

NOTE

Nr must be above 92% for RPM warning
light to be extinguished.

ABC 12. RPM caution system — Check that the RPM caution light will illuminate when:

GAS PROD (Ng) of either engine decreases to 52.5 ±.2% or ROTOR RPM decreases
to 92 ±2% Nr. The RPM caution audio signal comes on when rotor rpm decreases to
92 ± 2%.

3-43

Section III
Part 5

NAVAIR 01 -HIAAB-1

13. Boost pump check — Pull FUEL FWD BOOST pump circuit breaker. Check FWD
FUEL BOOST and MASTER CAUTION lights illuminate. Record aft boost pump
pressure (5-25 psi). Pull FUEL AFT BOOST pump circuit breaker. Check FUEL
AFT BOOST and MASTER CAUTION light illuminates and ensure that FUEL
PRESS goes to zero. Let the engines run for one minute to ensure that both engine
fuel pumps can draw fuel from the fuel cells. Reset FUEL FWD BOOST pump
circuit breaker and record forward boost pump pressure. Position the CROSS
FEED switch to CLOSED position. Check that the pressure goes to zero. Let the
engines run for 1 minute and reset FUEL AFT BOOST circuit breaker in. Turn
CROSS FEED ON. Check fuel pressure.

14. Generators:

Condition 1.

NON-ESS BUS - NORMAL
GEN 1 - ON
GEN 2 - OFF

Check No. 2 DC GEN caution light and MASTER CAUTION light illuminated;
AMPS 2 load zero. Non-Ess Bus electrical equipment operative. Record No. 1
DC GEN volts and amps. Repeat for No. 2 DC GEN.

Condition 2.

NON-ESS BUS - NORMAL
GEN 1 - OFF
GEN 2 — OFF

Check NO. 1 DC GEN and NO. 2 DC GEN caution lights and MASTER
CAUTION light illuminated; generator loads zero; all Non-Essential Bus
electrical equipment inoperative with AC MAIN and ALT ENCODER caution
lights illuminated. ALT ENCODER light will not function if AAU-32/A altimeter
is installed.

NON-ESS BUS switch — MANUAL. Ensure that Non-Essential electrical
equipment operative with battery power only.

Condition 3.

NON-ESS BUS - NORMAL
GEN 1 - ON
GEN 2 - ON

WEAPON CONT cb - OUTBOARD

3-44

MASTER ARM — STBY

Check NO. 1 DC GEN has assumed helicopter’s normal electrical load and NO. 2
DC GEN has assumed the armament system electrical load.

NAVAIR 01 -HI AAB-1

Section III
Part 5

ABC

Condition 4.

NON-ESS BUS - NORMAL
GEN 1 - ON
GEN 2 - OFF

WEAPON CONT cb - OUTBOARD
MASTER ARM - STBY

Check NO. 1 DC GEN has assumed the armament system electrical load.
Helicopter normal electrical load is on battery power with the Non-Essential
electrical equipment inoperative. The MASTER CAUTION light, NO. 2 DC GEN
caution light AC MAIN caution light, and ALT ENCODER caution lights on.

15. Inverters:

a. Main inverter — all electrical equipment on, 113.5 ±5 volts.

b. Standby inverter — all electrical equipment on, 113.5 ± 5 volts.

16. SCAS — Approximately 30 seconds after engaging SCAS POWER switch, the
NO-GO lights should be out provided SCAS channels are disengaged and controls
are stationary. Move cyclic forward slightly and note that NO-GO light
illuminates. Hold all controls stationary and note the PITCH light goes out within
30 seconds. Move cyclic back to center and note PITCH NO-GO light illuminates.
When NO-GO light goes out, engage PITCH channel. Repeat above for ROLL and
YAW using appropriate control movements. After all channels are engaged, check
cyclic SCAS release button in both cockpits. Move tip path twelve inches in pitch
and note SCAS corrects back; repeat for roll.

NOTE

Visually observe main rotor tip path
plane for excessive fluctuation when
ROLL and PITCH channels are
engaged. Excessive fluctuation is cause
to abort check.

17. Ecu, rain removal, and pitot heat check — Engergize ECU and check INLET
TEMP rise on both engine INLET TEMP gages (possible 10-25 degrees). Turn
ECU off. Energize RAIN RMV and note INLET TEMP rise and windshield.
Energize PITOT HTR and observe ammeter.

18. Compare cockpit instruments/gage. Record gage readings and note any gage
splits between cockpits in excess of:

a. 1% ROTOR RPM (Nr), ENG RPM (Nf) and GAS PROD (Ng).

b. 2% TORQUE

c. 10 degrees INLET TEMP.

19. UHF radio — Check for operation.

3-45

Section III
Part 5

NAVAIR 01 -HIAAB-1

ABC

20. Radar altimeter AN/APN-171(v) — ON. Allow 3 minutes warm-up, indicator OFF
flag disappear and altitude pointer indicates zero feet. Press, PUSH-TO-TEST, a
reading of 100 plus or minus 15 feet will be indicated if the system is functioning
properly.

Hover Checks.

1. Controls — Check helicopter performs correctly to control inputs by hovering
forward, rearward, sideward and turning left and right 360 degrees.

2. Pylon rock — With SC AS on, move cyclic fore and aft rapidly once or twice and
center cyclic. Induced oscillations should dampen out within 4-5 cycles. If
oscillations do not dampen, turn PITCH and ROLL SCAS OFF and repeat pro- II
cedure; if oscillations dampen normally after 4-5 cycles, the pitch SCAS may be “
defective.

3. Manual fuel — Set helicopter on deck and retard both throttles to engine idle. Place
ENGINE 1 GOV to MANUAL and cautiously roll No. 2throttlefull open. Roll No.

1 throttle on to join torque needles. Lift helicopter to a hover. Set helicopter back
down and return GOV to AUTO with throttles at engine idle. Repeat procedure for
No. 2 engine. Check ENG GOV MNL advisory lights illuminates. While GOV is in
MANUAL mode, check that the high rotor RPM caution light comes on at 103%
± 2 % ENG RPM (Nf).

4. SCAS yaw check—Once established in a stable hover, with the force trim ON, pull
in a 10-15% above hover torque without directional control input. Note that the
helicopter attempts to slow the yaw rate.

Flight Checks.

1. Torque limitor/droop cam check — At flat pitch, set ENG RPM (Nf) at 100%.
Smoothly execute a full power climb not to exceed limits: 100% XMSN
TORQUE, 101.8% GAS POD (Ng), or 837 degrees INLET TEMP. Check that
ENG RPM (Nf) does not vary by more than 4.5 — 5.5% momentarily then
restabilizes to 100% ± 1% ENG RPM (Nf). If Nf droops before engine limits are
attained, adjustments are necessary.

2. Power available:

a. Altimeter set at 29.92.

b. ECU and RAIN RMV OFF.

c. Climb to selected altitude and stabilize for one minute.

d. Reduce one throttle to engine idle. Increase test engine to full RPM increase
without changing engine trim.

3-46

NAVAIR 01 -HIAAB-1

Section III
Part 5

ABC

e. Without exceeding helicopter or engine limits, increase power on the test engine
until ENG RPM (Nf) droops to 97% or until an INLET TEMP or GAS PROD
(Ng) limit is reached. Allow inlet TEMP to stabilize for a minimum of 30
seconds up to a maximum of 3 minutes, stabilizing only long enough to satisfy
requirements. If INLET TEMP indicates 837°C prior to obtaining the ENG
RPM (Nf) droop, climb to a higher altitude and repeat the test. If droop does
not occur at 101.8 ± .5% GAS PROD (Ng) engine adjustments are necessary.

NOTE

At extremely low OAT, ENG RPM (Nf)
droop may occur prior to reaching
INLET TEMP or GAS PROD (Ng)
limits. This is a result of the engine
reaching a maximum fuel flow limit of
640-660 pph. In addition, certain engines
will not reach 101.8% GAS PROD (Ng)
regardless of altitude, without exceeding
the maximum INLET TEMP limit due to
their internal speed and temperature
match. This does not prevent the engine
from achieving required performance.

In these cases, record normal topping
parameters at 97% ENG RPM (Nf) (beep
if necessary, but maintain an INLET
TEMP of 837°C by changing collective
position) and compare them to the AH-
1T topping chart requirements to insure
that proper engine power is available. In
order to set Maximum GAS PROD (Ng)
limit stop for an INLET TEMP limited
engine, utilize the part power trim stop
and set GAS PROD (Ng) to 98.3 to
98.8%.

f. Insure that ENG RPM (Nf) is stabilized at 97%, then record OAT, altitude,
ENG RPM (Nf), GAS PROD (Ng), INLET TEMP, and TORQUE.

g. Verify that test torque equals or exceeds the requirements of the power avail¬
able chart. If chart torque cannot be equalled, instrument calibration and/or
engine maintenance is required.

NOTE

Chart torque values have been reduced
to allow for the most probable low
readings due to allowable
instrumentation error.

3. Hydraulic boost checks — In stabilized level flight at 100 KIAS, turn off the No. 2
hydraulic system, being prepared to return to ON if control forces become
excessive in any manner. Note the following:

a. NO. 2 HYD PRESS caution light ON

Change 1 3.47

Section III
Part 5

ABC

NAVAIR 01-H1AAB-1

1). MASTER CAUTION light ON
c\ I1Y I) 2 PSI indication near zero

d. PITCH and ROLL SCAS switches drop to “OFF”

e. Loss of pitch and roll SCAS

Perform .‘JO degree angle of bank turns in both directions and moderate nose up
and nose down maneuver.

f. The NO. 1 hydraulic system should provide normal boost-on control responses
in all controls without feedback, rate limiting, or motoring.

HY1) switch ON and note normal operation. Engage PITCH and ROLL SCAS.
Turn off Hydraulic 1 system with same precaution as above. Note the following:

a. NO. 1 HY1) PRESS caution light ON

b. MASTER CAUTION light ON

c. HYI) 1 PSI indication near zero

d. YAW SCAS switch drops to OFF

e. Loss of yaw SCAS

1. 1 he NO. 2 hydraulic system should provide normal boost-on control responses
in the cyclic and collective controls without feedback, rate limiting, or
motoring. The directional pedals will be non-boosted but should be manageable
through all normal maneuvers including hover and landing.

HYI) switch ON and note normal operation. Engage YAW SCAS.

4. Normal Maneuvers — At airspeeds above 60 knots, the right pedal should be
slightly ahead of the left pedal. 30-degree angle of bank left turns, at 55 knots,
should be possible at full power. 30-degree angle of bank right turns should be
possible while autorotating at 100 knots. There should be no need for excessive
lateral control as airspeed is increased.

5. 1 racking and vibration — A vibration analyzer unit should be utilized to evaluate
excessive helicopter vibrations. Refer to the appropriate maintenance manual for
installation of unit, flight procedures, and data analysis.

6. Autorotation rpm check — Autorotation to be conducted at 70 KIAS, balanced,
unaccelerated, wings level flight, with collective full down. Record OAT altitude
gross weight, and ROTOR RPM (Nr).

7. Avionics — Check operation of all equipment as stated in the appropriate
technical manual.

3-48

ROTOR RPM

ABC

NAVAIR 01-H1AAB-1

AT HEAVY GROSS WEIGHT OR HIGH ALTITUDE,
SLIGHTLY UP COLLECTIVE MUST BE USED TO
PREVENT ROTOR OVERSPEED.

Figure 3-9. Rotor Limit Chart

Shutdown.

1. Engine shutdown — In accordance with Checklist, Section III, Part 2.

2. Check PSI following caution lights illuminate out:

ENG

PSI

XMSN

PSI

C BOX

PSI

3. Rotor brake light.

Section III
Part 5

210900-73

3-49/(3-50 blank)

NAVAIR 01-H1AAB-1

Section IV

SECTION IV— FLIGHT CHARACTERISTICS

TABLE OF CONTENTS

Introduction.4-1

Rotor Blade Stall .4-1

Control Feedback .4-1

Pitch — Cone Coupling.4-2

Maneuvering Flight.4-2

Radius of Turn.4-2

Low "G” Maneuvers.4-2

Mast Bumping.4-3

Diving Flight.4-3

Hovering Capability.4-6

Dynamic Rollover Characteristics.4-6

Pylon Rock.4-7

Power Settling.4-7

Rotor Droop .4-7

Vibration Identification .4-7

Autorotation Characteristics.4-8

INTRODUCTION.

The flight characteristics of this helicopter are
similar to other single rotor helicopters. The basic
flying qualities are enhanced by the Stability and
Control Augmentation System (SCAS). This system
provides good stability and control response
throughout the operating flight envelope. The
control system, with hydraulic servo assist, provides
the pilot with a light force required for control
movements; control feel is induced into the cyclic
stick and tail rotor controls by means of a force trim
system.

ROTOR BLADE STALL.

NOTE

Main rotor blade stall is not a problem in
the AH-IT (TOW) helicopter when
operated within the approved flight
envelope. However, main rotor blade
stall may occur at some combination of
excessive airspeed and high “G”
loading.

Blade stall occurs when the angle of attack of the
retreating blade exceeds the specific stall angle for
any blade segment. When the condition is attained,
increased blade pitch (or collective) will not result in
increased lift and may result in reduced lift. The
threshold of stall is approached as gross weight,
airspeed, altitude and "G” loading increase and rpm
decreases. One of the more important features of the
two-bladed, semi-rigid system is its warning to the
pilot of impending blade stall. Prior to progressing
fully into the stall region, the pilot will feel a marked
increase in airframe vibration and, possibly, control
feedback. Consequently, corrective action can be
taken before stall becomes severe.

Blade Stall — Corrective Action.

The use of the following procedures is predicated
on the helicopter’s altitude above the terrain.
Sufficient recovery altitude must be available for
these to be effective. When blade stall is evident the
condition may be eliminated by accomplishing
one or a combination of the following corrective
actions.

1. Reduce collective.

2. Reduce airspeed.

3. Decrease severity of maneuvers.

4. Increase operating rpm.

5. Descend to lower altitude, if appropriate.

CONTROL FEEDBACK.

Feedback in the cyclic stick or collective stick
caused by high loads in the control system. These
loads are generated during severe maneuvers and
can be of sufficient magnitude to overpower or feed
through the main boost cylinders and into the cyclic
and/or collective stick. The pilot will feel the feedback
as an oscillatory "shaking” of the controls even
though he may not be making control inputs after the
maneuver is established. This type of feedback will
normally vary with the severity of the maneuver.
The pilot should regard it as a cue that high control
system loads are occurring and should immediately
reduce the severity of the maneuver.

4-1

Section IV

NAVAIR 01 -HIAAB-1

The copilot/gunner station side arm flight
controls have a reduced mechanical
advantage. Because of this reduced
mechanical advantage of the
copilot/gunner cyclic and collective
controls, severe maneuvers should be
avoided while flying from the gunner
station. If the pilot-in-command elects to
allow maneuvers to be flown from the
copilot/gunner station, the rear seat pilot
should monitor the flight controls and be
capable of recovering to a safe attitude if
required.

PITCH-CONE COUPLING.

Pitch-cone coupling is the tendency of the rotor
blade to reduce pitch as thrust is increased or rotor
RPM is reduced. With large amounts of pitch-cone
coupling, the rotor may overspeed during pull-ups
or flares unless the pilot adds collective pitch. The
AH-1T (TOW) main rotor design minimizes pitch
cone coupling.

MANEUVERING FLIGHT.

When performing maneuvers above 120 KIAS, it is
necessary to devote more attention to flying and to
planning manuevers due to the increased distance
needed to perform pull outs and turns. The increased
distance required for pull outs and turns is a direct
result of the higher airspeed.

CAUTION

During left ‘rolling maneuvers or high
power dives, torque, Ng and ITT
increases occur. Care shall be exercised
to monitor instruments, especially the
triple tachometer. This will enable the
pilot to adjust power as required to
prevent exceeding aircraft engine
limitations and prevent a low rotor
RPM condition. This can be accomplished
by either lowering the collective, reduc¬
ing the severity of the maneuver, or a
combination of both.

Radius Of Turn.

At airspeeds above 130 KIAS the radius of turn and
rate of closure increases rapidly due primarily to
higher airspeeds. The turn radius is a function of the
bank angle ("G” loading) and the square of the
airspeed. For any given condition of altitude and
weight, where the "G” capability is defined by rotor
characteristics the turn radius can be markedly

affected by airspeed. The effect of speed can be
ascertained by an inspection of figure 4-1. From the
examples A and B, it can be seen that for a bank angle
of 30° (1.15 "G”) the radius of turn is increased by a
factor of four when the airspeed is increased from 80
KIAS to 160 KIAS. The same is also true in a dive
recovery. Figure 4-2 provides a graphic chart of the
turning radius in relationship to airspeed.

Low "G" Maneuvers.

AH-1T (TOW) helicopters have a tendency to roll
to the right when forward cyclic is used to initiate
a lower than one G maneuver in forward flight.
The reason for this low G roll tendency is the
thrust produced by the tail rotor. Because the tail
rotor is above the helicopter’s center of gravity,
the tail rotor thrust produces a right roll tendency.
During normal one G flight, a portion of the main
rotor thrust balances the tail rotor thrust and
counteracts this right roll tendency. During low G
flight, however, main rotor thrust is greatly
reduced while the tail rotor thrust remains high,
thus, a right roll can develop during low G maneu¬
vers. Instinctive pilot reaction is to correct the roll
with left lateral cyclic. But since main rotor thrust
has been greatly reduced, lateral cyclic effective¬
ness is also greatly reduced. Left cyclic application
may also result in mast bumping. Aft cyclic will
quickly increase rotor thrust (higher G) and will
return lateral cyclic effectiveness.

Because of mission requirements, it may be neces¬
sary to rapidly lower the nose to (1) acquire a
target, (2) stay on target or, (3) recover from a
pullup. At moderate to high airspeeds, fairly small
abrupt forward cyclic inputs can yield G levels near
zero. The helicopter may roll to the right simul¬
taneously with forward cyclic, the roll being
greater as G levels approach zero and when the
roll SCAS is disengaged. If an abrupt right roll
should occur when rapidly lowering the nose,
pull in aft cyclic to stop the roll and effect
recovery. Left lateral cyclic will not effect recovery
from a well developed right roll during the flight
below 1 G and may cause mast bumping. Do not
engage/disengage SCAS during recovery. When it
is necessary to rapidly lower the nose, it is
essential that the pilot monitor changes in roll
attitude as the cyclic is moved forward.

Should an uncommanded right roll occur during
flight below one G, the following procedures are
recommended:

1. Cyclic — Immediately center lateral, then aft.

4-2

Change 1

NAVAIR 01-H1AAB-1

Section IV

When main rotor returns to normal thrust
conditions:

2. Controls — As required to regain balanced
flight.

If mast bumping occurred or was suspected:

3. Land as soon as possible.

Should flight conditions occur in which
the above procedures are impractical
(i.e., nose extremely high at low air¬
speed), use pedals to yaw aircraft into
nose low dive. Up collective can also be
used when power available permits. Do
not engage/disengage SC AS during
recovery.

Mast Bumping.

Mast bumping occurs when the rotor exceeds its
critical flapping angle and the underside of the
rotor hub contacts (bumps) the rotor mast. If
contact is severe, mast deformation can occur and
cause mast structural failure. Excessive rotor
flapping can also cause rotor blade contact with
the tailboom or cockpit. Mast bumping generally
occurs at, but is not restricted to, the extremes of
the operating envelope. The most influential
causes are (in order of importance):

1. Low G maneuvers (below plus 0.5 G).

2. Abrupt roll reversals (larger flapping occurs
during left to right reversals).

3. Rapid large cyclic motion (especially
forward cyclic).

4. Flight near longitudinal/lateral CG limits.

5- High slope landings.

Less significant causes are maximum sideward/
rearward flight, sideslip, and bladestall conditions.

WARNING

Should mast bumping occur in flight,
catastrophic results are highly probable.
Since conditions causing rotor flapping
are cumulative, improper pilot response/
recovery techniques to flight situations
approaching or favorable to mast bump¬
ing can aggravate the situation and lead
to in-flight mast bumping and mast
separation.

Favorable conditions and recommended recovery
procedures for mast bumping are provided below.

CONDITION

RECOVERY TECHNIQUE

Start/Shutdown

Cyclic: Move to stop bumping.

Rear/Side Flight

Cyclic: Move slightly toward
center.

Pedal: Bring nose into wind.

Slope Landing

Cyclic: Move toward center to
stop bumping, re¬
establish hover.

Engine failure at
high forward
airspeed

Cyclic: Move aft to maintain
positive G (positive
thrust), retain Nr and
avoid mast bumping
during auto entry.

Collective: As req’d to main¬
tain Nr.

Low G maneuvers
(below plus

0.5 G) (other than
nose high)

Cyclic: Center laterally and aft
to regain positive G
(positive thrust) on the
rotor & maintain Nr.

Nose high, low
airspeed

Collective: Judiciously in¬

crease, if possible.

Pedal: As req’d to establish
nose low condition.
Cyclic: Neutral.

Diving Flight.

Diving flight presents no particular problems in
the AH-IT (TOW); however, the pilot should have a
good understanding of such things as rates of
descent versus airspeed, rate of closure and rates of
descent versus power. Because of relatively low
drag, the helicopter gains airspeed quite rapidly in
a dive and it is fairly easy to exceed the redline.
Rates of descent over 3000 ft./min. are not
uncommon during high speed dives. These high
rates of descent coupled with the high flight path
speeds (290 ft./sec. at 170 KIAS) require that the
pilot monitor both rate of closure and terrain
features very closely and plan his dive recovery in
time to avoid having to make an abrupt recovery.
If an abrupt recovery is attempted at speeds near
redline airspeed, “mushing” of the helicopter can
occur. If mushing is experienced, do not increase
collective. Application of increased collective will
aggravate condition.

Change 1 4-3

Section IV

NAVAIR 01-H1AAB-1

TURN RADIUS= -

g ton ct>

NORMAL LOAD FACTOR^ —!—

COS <J)

TURN RADIUS

Note This chart gives the turn radius in feet as a function of
airspeed and either bank angle or normal load factor.
The capability of the aircraft is not inferred by this
chart, but trade-off of bank angle versus turn radius
are valid.

‘G’ LOADS

EXAMPLE A

AIRSPEED - 80 KTAS
BANK ANGLE - 30 DEGREES

SOLUTION:

TURN RADIUS - 981 FEET
‘G' LOAD - 1.15

EXAMPLE B

AIRSPEED - 160 KT AS
BANK ANGLE - 30 DEGREES

SOLUTION:

TURN RADIUS - 3925 FEET
‘G’ LOAD - 1.15

20«>«>00-31

Figure 4-1. Radius of Turn

4-4

FEET TURN RADIUS - FEET

NAVAIR 01-H1AAB-1

Section IV

Note

This chart gives the turn radius in feet as a function of airspeed
and either bank angle or normal load factor. The capability of
the aircraft is not inferred by this chart, but trade-off of bank
angle versus turn radius are valid.

209900-30

Figure 4-2. Radius of Turn — 30 Degree Bank

4-5

Section IV

NAVAIR 01 -HI AAB-1

POWER DIVES.

At speeds above the maximum level flight speed, the
rate of descent will increase approximately 1000
ft./min. for every 10 knots increase in airspeed for the
full power condition.

HOVERING CAPABILITY.

Hovering capability is affected by in-ground-effect
(IGE), out-of-ground effect (OGE), outside air
temperature (OAT), pressure altitude, wind speed,
engine torque (power available), and gross weight
of the helicopter. Hovering IGE performance is
better than OGE because during IGE the rotor sets
up a current flow between the helicopter and the
ground, providing a cushion of air to partially
support the helicopter weight. Temperature
variations affect engine and rotor performance.
Hovering with heavier gross weights or at higher
altitudes is possible with lower temperatures and
higher wind velocities. Lower temperatures
increase engine efficiency and wind represents
airspeed; therefore, either condition or both, will
increase hovering performance due to the ability of
the main rotor to provide more lift.

DYNAMIC ROLLOVER
CHARACTERISTICS.

During normal takeoffs and landings, slope
takeoffs, and landings, or landings and take-offs
with some bank angle or side drift, the bank angle
or side drift can cause the helicopter to get into the
situation where it is pivoting about a skid. When
this happens, lateral cyclic control response is
more sluggish and less effective than for the free
hovering helicopter. Consequently, if the bank
angle (the angle between the aircraft and the
horizon) is allowed to build up past 15°, the
helicopter will enter a rolling maneuver that
cannot be corrected with a full cyclic and the
helicopter will roll over on its side. In addition, as
the roll rate and acceleration of the rolling motion
increases, the angle at which recovery is still
possible is significantly reduced. The critical roll
over angle is also reduced for a right skid down
condition, crosswinds, lateral center of gravity
offset and left rudder pedals inputs.

When performing maneuvers with one skid on the
ground, care must be taken to keep the aircraft
trimmed, especially laterally. For example, if a
slow takeoff is attempted and the tail rotor thrust
contribution to rolling moment is not trimmed out
with cyclic, the critical recovery angle will be

exceeded in less than 2 seconds. Control can be
maintained if the pilot maintains trim, does not
allow aircraft rates to become large, and keeps the
bank angle from getting too large. The pilot must
fly the aircraft into the air smoothly keeping
executions in pitch, roll and yaw low and not
allowing any untrimmed moments.

When performing slope take-off and landing
maneuvers, follow the published procedures, being
careful to keep roll rates small. Slowly raise the
down slope skid to bring the aircraft level and then
lift off. (If landing, land on one skid and slowly
lower the down slope skid). If the aircraft rolls to
the up slope side (5° to 8°), reduce collective to
correct the bank angle and return to wings level
and then start the take-off procedure again.

Collective is much more effective in controlling the
rolling motion than lateral cyclic because it
reduces the main rotor thrust. A smooth, moderate
collective reduction of less than approximately
40% (at a rate less than approximately full up to
full down in 2 seconds) is adequate to stop the
rolling motion with about 2 degrees bank angle
overshoot from where down collective is applied.
Care must be taken to not dump collective at too
high a rate as to cause fuselage-rotor blade
contact. Additionally, if the helicopter is on a slope
and the roll starts to the up slope side, reducing
collective too fast creates a high rate in the
opposite direction. When the low slope skid hits the
ground, the dynamics of the motion can cause the
aircraft to roll about down slope skid and over on
its side. Do not pull collective suddenly to get
airborne as a large and abrupt rolling moment in
the opposite direction will result. This moment
may be uncontrollable.

If the aircraft reaches 15° of bank angle
with one skid on the ground and thrust
approximately equal to the weight, the
aircraft will roll over on its side. Reduce
collective to stop the roll and correct the
bank angle to wings level.

CAUTION

VV^hen landing or taking off, with thrust
approximately equal to the weight and
one skid on the ground, keep the aircraft
trimmed and do not allow aircraft rates
to build up. Fly the aircraft smoothly off
(or onto) the ground, carefully
maintaining trim.

4-6

NAVAIR 01-H1AAB-1

Section IV

PYLON ROCK.

The AH-1T (TOW) is not subject to pylon rock
under normal conditions. Pylon rock is the
phenomenon of the helicopter pylon moving
periodically ( V ,2 per Rev or 2.7 cps). This pylon
motion is commonly noted by several short self¬
damping oscillations with the number of
perceptible oscillations indicative of the state of
wear of pylon dampers.

If pylon rock is encountered, a change of flight
condition, preferably by lowering the collective,
should eliminate the motion.

POWER SETTLING.

Power Settling is most likely to occur during
conditions of high gross weight, high density
altitude, low airspeed and descending powered flight.
Under these conditions a helicopter is settling
through the air displaced by its own rotor system.
The downwash then recirculates through the
helicopter rotor system, resulting in reduction of
lift, increased roughness and poor control response.

Power settling is an uncommanded rate of descent
caused by the helicopter rotor encountering the
“vortex ring state” as it settles into its own down-
wash. In this state the flow through the rotor
system is upward near the center of the rotor disc
and downward in the outer portion. This results
in zero net thrust from the rotor and extremely
high aircraft descent rates. Power settling is not
restricted to high gross weights or high density
altitudes. It may not be recognized, and a
recovery effected, until considerable altitude
has been lost. Helicopter rotor theory indicates
that it is most likely to occur when descent rates
exceed 800 FMP during (1) vertical descents
initiated from a hover and (2) steep approaches
at less than 40 knots.

Indications to the pilot are:

1. Rapid descent rate increase.

2. Increase in overall vibration level.

3. Loss of control effectiveness.

Recovery by:

1. Forward cyclic to gain airspeed.

2. Descrease collective.

Increasing collective has no effect
toward recovery and will aggravate
power settling. During approaches at
less than 40 knots, avoid descent rates
exceeding 800 FPM.

ROTOR DROOP.

Droop is a term used to denote a change in power
turbine speed (Nf) and rotor speed that occurs with a
demand for increased power with the governor at a
constant speed setting. Droop may be further
categorized as either transient or steady state.
Transient droop is the momentary change in power
turbine speed and rotor speed resulting from an
increased power demand, and it is compensated for by
the Nf governor. Steady state droop is the decrease in
power turbine speed and rotor speed which results
from an increased power demand (stabilized
condition) and it is not compensated for by the Nf
governor control.

VIBRATION IDENTIFICATION.

One/Rev Vibration (Main Rotor).

This vibration is relatively easy to recognize in that it
is quite easy to count (approximately 5/sec.). The
following are normal causes of 1/rev vibration:

1. Rotor out of balance condition causes a lateral
1/rev. vibration. The rotor can be out of balance
either chordwise or spanwise. An
out-of-balance condition can appear as a
vertical vibration during forward flight.
Therefore, it is best to balance the rotor before
attempting to analyze other 1/rev. vibrations.
Consult the MIMs for corrective action.

2. Rotor out of track condition causes a vertical
vibration and will normally increase in
amplitude with airspeed. Appropriate
corrective action is outlined in the MIMs.

3. Binding in the scissor links or mixing levers.

4. Binding in rotor grip bearings.

Low Frequency Vibration (Pylon Rock).

This vibration manifests itself as a vertical vibration
(about 3/sec). It is more noticeable at low airspeeds
and high power, at forward eg. This "Rocking Chair”
motion can usually be reduced by reducing speed and

4-7

Section IV

NAVAIR 01 -HI AAB-1

power. It is the result of the pylon mounts either
having failed or deteriorated. It can also be induced
by erratic cyclic motion.

Two/Rev Vibration.

This vibration (10/sec) is extremely difficult to count.
Amplitude increases with airspeed as a result of
unequal drag causing the top of the mast to move in a
manner to shake the pylon at 2/rev frequency. This
can be caused by soft pylon mounts, although a
certain amount of 2/rev is inherent in the helicopter.

1. Check pylon mounts for separation or
bottoming out.

2. Check drag braces of the rotor to see that they
are mounted securely and have no play in
attachment points.

3. Tailboom attachment bolts.

High Frequency Vibration.

High frequency vibrations are much too fast to count
and feel like a "buzz”. These frequencies may
emanate from the engine, improper driveshaft
alignment, couplings improperly functioning,
bearings dry or excessively worn, or tail rotor out of
track or balance. If excessive high frequency
vibration exists, it is recommended that the
helicopter land and a crew member attempt to locate
the source. The area where the highest amplitude of
the vibration exists is generally the area from which
the vibration is originating.

AUTOROTATION CHARACTERISTICS

Due to the wide speed range capability of the AH-
1T (TOW), some discussion of the POWER OFF
characteristics of the rotor system is essential.

Main Rotor.

The following steps explain the necessity of
maintaining rotor rpm in its normal power off range
(91 to 105 percent).

Normal Rotor Speed.

The normal rotor speed assures the pilot that he will
retain adequate control effectiveness. Low rpm
(underspeed) causes a proportional loss of response to
control inputs. High rpm (overspeed) can cause
structural damage to the rotor system.

Rotor Flapping.

The angle between the tip path plane and the mast
increases at low rpm. By maintaining rotor rpm in
the normal range, the pilot assures safe clearance
between the rotor and the tailboom.

Rotor Inertia.

Rotor inertia is a characteristic which tends to
prolong the effectiveness of collective control in the
autorotation landing. This effectiveness decreases
with rpm. Normal rotor rpm assures the pilot that he
will have normal inertia and normal collective
control response with which to arrest the sink rate in
the autorotation landing.

Rotor RPM.

The following steps list the factors which affect
power-off rotor rpm.

AIRSPEED.

In autorotation, rotor rpm varies with airspeed.
Maximum rotor rpm is achieved at a steady state
of 60 to 80 KIAS (Figure 4-3). Rotor rpm decreases
at stabilized airspeeds above or below 60 to 80
KIAS range. When changing airspeeds, cyclic
movement will produce a rotor rpm other than that
produced under steady state conditions as follows:

FROM LOW AIRSPEED. Example: From a
stabilized 30 KIAS autorotative condition, a positive
forward cyclic movement to increase airspeed will
cause the rotor rpm to decrease initially and then
increase when the helicopter is stabilized at the
higher speed.

FROM HIGH AIRSPEED. Example: From a
stabilized 120 KIAS autorotative condition, a
positive aft cyclic movement to decrease airspeed will
cause the rotor rpm to increase initially and decrease
when the helicopter is stabilized at the lower speed.

4-8

Change 1

NAVAIR 01-H1AAB-1

Section IV

110%

5 100%

90%

80%

- (

:oi

NS'

fAI

‘>S\

/VE

IGF

HT,

\LTITUDE AND COLLECTIV
POSITION

r E

1 1 1 1 1 1 1 1 1 1 1 1

HIS CHART IS AN EXAMPLE
<ND IS USED ONLY TO EXPLAIN
)ATA IN THIS SECTION

- 7

M 11 11 II 11 11

20 40 60

80 100 120
V CAL KNOTS

140 160 180 200

210900-77A

Figure 4-3. Autorotation RPM Versus Airspeed

NOTE

The maximum permissible steady state
autorotation airspeed is 120 KIAS.

GROSS WEIGHT.

The power-off rpm varies significantly with gross
weight for identical collective settings. A low gross
weight will produce a low rotor rpm. A high gross
weight will produce a high rotor rpm. With the
collective system correctly rigged to a minimum
blade angle (full down collective stick) of
approximately 6.75 degrees the pilot must manually
control rpm with collective stick in order to prevent
overspeeding of the rotor when at high gross weight.

DENSITY ALTITUDE.

The power-off rotor rpm varies with altitude; low
altitude — low rpm; high altitude — high rpm. The

pilot will find that the higher the altitude — the
higher the collective stick position required to
prevent overspeed of the rotor.

CYCLIC FLARE.

Aft cyclic control application (nose up pitching)
produces an increase in rotor rpm proportional to the
flare and entry speed. The higher the speed — greater
the flare effectiveness. From a high speed entry
condition, a steep flare can produce an overspeed
unless limited by collective pitch control.

Pilot Technique.

It can be readily seen from the information, that the
pilot technique must vary in accordance with the
actual conditions of airspeed, altitude, and gross
weight at the time of engine failure.

4-9/(4-10 blank)

SECTION V — EMERGENCY PROCEDURES

TABLE OF CONTENTS

Introduction.5-1

Advisory Caution and Warning Light —

Initial Action.5-2

PART 1 - GROUND EMERGENCIES

Emergency Egress and Rescue.5-5

Hot Start.5-7

Engine Fire on Start (External).5-7

PART 2 - TAKEOFF EMERGENCIES

Single Engine Failure During Takeoff.5-7

Dual Engine Failure During Takeoff.5-8

PART 3 - INFLIGHT EMERGENCIES

Hydraulic Malfunctions. 5-8

SCAS Failure.5-11

Control System Malfunctions.5-12

Tail Rotor Malfunction.5-12

Mast Bumping.5-13

Uncommanded Right Roll

During Flight Below 1 G.5-14

Engine Malfunctions.5-14

Main Driveshaft Failure.5-21

Electrical System Malfunctions.5-21

Elimination of Smoke and Fumes

in Cockpit.5-23

Electrical Fire.5-24

Fuselage Fire In Flight.5-24

Fuel System Malfunctions.5-25

Impending Transmission Failures.5-25

Combining Gearbox Malfunctions.5-26

42° and 90° Gearbox Malfunctions.5-27

Rotor Brake Pressurized In Flight.5-28

Wing Stores Jettison.5-28

Lost Plane Procedures.5-28

Lost Sight During IMC.5-28

PART 4 - LANDING EMERGENCIES

Autorotative Landing.5-30

Single Engine Landing.5-30

Landing in Trees.5-31

Ditching.5-31

INTRODUCTION.

Emergency Procedures.

actions are those that contribute to an orderly
sequence of events and assure that all necessary
actions are taken. These procedures are
accomplished with direct reference to the check¬
list.

Emergency procedures are divided into two
categories, critical and non-critical. The critical
items are those which must be performed
immediately if the emergency is not to
be aggravated. These critical items are underlined
and must be performed immediately in proper
sequence. Non-critical emergency procedure

Scope.

The following procedures contain the indications
of failures or malfunctions which affect safety of
the crew, the helicopter, ground personnel or
property; the use of emergency features of primary
and back-up systems; and appropriate warnings,
cautions, and explanatory notes.

Special Instructions.

1. The following terms indicate the degree of
urgency in landing the helicopter.

Land immediately

Self-explanatory. Landing
in trees, water or otherwise
unsafe areas should be
considered as a last resort.

Land as soon as possible - Land at first site of which

a safe landing is reasonably
assured.

Change 1

5-1

Section V

NAVAIR 01-H1AAB-1

Land as soon as practical - Extended flight is not recommended,

and landing site and duration

of flight are at the discretion of the pilot.

2. The following terms are used to describe the
operating condition of a system, subsystem,
assembly or component.

Affected

Normal

- fails to operate in the normal
or usual manner.

- operates in the normal or usual
manner.

3. The master caution light will illuminate
when any caution panel light illuminates.

When caution and warning lights are
illuminated, accomplish the actions and
procedures as follows:

*Only on pilot panel
**Segments aviation green

ADVISORY CAUTION AND WARNING LIGHT - INITIAL ACTION

PANEL WORDING

CONDITION

CORRECTIVE ACTION

OIL PRESS
.(engine #1,

1 engine #2)

Respective engine oil
pressure below operat¬
ing minimum (40 psi).

Check engine oil pressure
to verify low indication.

If below limit, shut down
respective engine. Land
as soon as practical.

CHIP DETR

(engine #1,
engine #2)

Metal particles in

respective engine.

Flight idle. Check oil S

pressure and temperature. “

If normal operate at
reduced power. If pressure is
low and/or temperature is
high, shut down respective
engine. Land as soon as
practical.

FUEL FILTER

(engine #1,
engine #2)

Fuel filter partially

obstructed.

Prepare for single engine
failure. Land as soon as
practical.

*PART SEP OFF

Particle separator

door not full open
(respective engine).

Check switch position —

OFF or AUTO. Verify engine
rpm. Avoid continued
operation in icing or engine
eroding environment.

NO. 1 — 2 DC GEN

Respective dc

generator failed.

GEN-RESET, then

ON. If DC GEN light
remains illuminated,

GEN — OFF.

NAVAIR 01 -HIAAB-1

Section V

ADVISORY CAUTION AND WARNING LIGHT — INITIAL ACTION (Cont d.)

PANEL WORDING

* 90°TEMP/PRESS

*42° TEMP/PRESS

XMSN CHIP DETR

FAULT CONDITION

CORRECTIVE ACTION

OIL TEMP above limits
and/or OIL PRESS below
operating minimum.

Land as soon as possible.

OIL TEMP above limits

and/or OIL PRESS below
operating minimum.

Land as soon as possible.

Particles in
transmission.

Power 75% or less. Land as

soon as possible. If XMSN OIL
HOT OR XMSN OIL PRESS
light illuminates, reduce power
to 60% or less. Refer to
Impending XMSN Failure.

C BOX CHIP DETR

Metal particles in
combining gearbox.

Reduce power. Land as
soon as possible.

90° CHIP DETR

Metal particles in

90° gearbox.

Reduce tail rotor power.

Land as soon as practical.

42° CHIP DETR

Metal particles in

42° gearbox.

Reduce tail rotor power.

Land as soon as practical.

XMSN OIL HOT

Oil overtemperature
(above 110°C).

Verify oil temperature. Reduce

power. Land as soon as possible.
Refer to Impending XMSN Failure.

XMSN OIL PRESS

Oil pressure below

operating minimum
(30 psi).

Verify oil pressure. Reduce power

to 60% or less. Land as soon as
possible. Refer to Impending
XMSN Failure.

C BOX OIL PRESS

Oil pressure below

operating minimum
(40 psi).

Verify oil pressure.

Reduce power. Land as
soon as possible.

C BOX OIL HOT

Oil overtemperature
(above 116°C).

Verify oil temperature.

Reduce power. Land as
soon as possible.

NO. 1 HYD PRESS

NO. 2 HYD PRESS

Respective hydraulic

system below operating
minimum (below 2200
psi).

Check pressure. If pressure

is low, shut off affected
system. Land as soon
as possible.

NO. 1 HYD TEMP

NO. 2 HYD TEMP

Respective system oil

overtemperature.

Shut off affected system.

Land as soon as possible.

Change 1 5-3

5-3 A

r /5

NAVAIR 01 -HIAAB-1

ADVISORY CAUTION AND WARNING LIGHT - INITIAL ACTION (Cont'd.

PANEL WORDING

CONDITION

CORRECTIVE ACTION

* BATTERY TEMP

Battery overheating.

Turn off BATTERY

Land as soon as possible.

* XMSN OIL BYP

Oil bypassing cooler.

Reduce power to 60% or less.

Land as soon as possible. Refer
to Impending XMSN Failure.

AC MAIN

Main inverter failure.

INVERTERS — MAIN, check
circuit breaker In. If
inverter is still inopera¬
tive, INVERTERS — STBY.

Turn off all non-critical
ac equipment.

AC STBY

Stand-by inverter

failure.

INVERTERS — STBY, check
circuit breaker In.

FWD-AFT FUEL LOW

Respective fuel cell

quanity low.

CROSSFEED - OPEN.

Land as soon as possible.

*FWD-AFT FUEL

BOOST

Fuel boost pump pressure

low (below 5 psi).

CROSSFEED — OPEN. Pull
respective circuit breakers.

**ENG 1-2 GOV MAN

Respective engine

governor operating
in manual mode.

Pilot controls engine rpm
with twist grip throttle.

*AMMO DOOR OPEN

Ammunition compartment

door open.

Close door.

*ALT ENCODER

Electrical power lost
to altimeter encoder.

If AAU-32/A not
installed.

None

*EXT PWR DOOR

OPEN

External power door

open.

Close door.

**IFF

KIT-1A T/SEC

ZEROIZED.

Check IFF switches
and circuit breaker.

FIRE 1 PULL

Fire in engine No. 1.

Pull handle/FIRE EXT switch.
MAIN. Shut down engine.

Select RESERVE if required.

Land as soon as possible.

FIRE 2 PULL

Fire in engine No. 2.

Pull handle/FIRE EXT switch.
MAIN. Shut down engine.

Select RESERVE if required.

Land as soon as possible.

NAVAIR 01-H1AAB-1

Section V
Part 1

ADVISORY CAUTION AND WARNING LIGHT — INITIAL ACTION (Cont'd.)

PANEL WORDING CONDITION CORRECTIVE ACTION

*ROTOR BRAKE Rotor brake engaged. Place rotor brake handle

down. If light remains
illuminated, shut off
hydraulic system # 2. If light
does not extinguish,
land as soon as possible.

MASTER CAUTION

Segment in caution
panel illuminated.

Check caution panel.

RPM

Rotor rpm high, low, or

Check triple tachometer

engine rpm low.

and correct rpm as

as required.

PART 1 — GROUND EMERGENCIES

EMERGENCY EGRESS AND RESCUE.

Pilot and copilot/gunner access is provided by
canopy doors that are hinged at the top and swing
outward and up. Both doors can be opened or
closed either manually or electrically from inside
or outside. Emergency exit or entrance is provided
by a det cord system to cut the windows from the
canopy support structure. The linear explosive
system is installed around both canopy doors and
around the windows on each side. Interconnecting
lines of flexible detonating cord connect the linear
explosive system with the three canopy jettison
handles. The canopy jettison handles are located
on the left side of the pilot glare shield, the
copilot/gunner right console, and one is installed
in the nose of the helicopter for ground rescue
personnel. The system can be actuated from any of
the canopy jettison handles (figure 5-1).

Emergency Egress.

1. Lap belt/shoulder harness release —OPEN.

2. Helmet - DISCONNECT (HSS/ICS).

3. Canopy door handle — ROTATE (upward).

4. Manual clutch release — PUSH.

5. Helicopter — EXIT.

If canopy door cannot be opened manually:

1. Canopy jettison handle — ROTATE (90
degrees counterclockwise) and PULL.

WARNING

Personnel positioned within 50 feet of
the helicopter could be injured by debris.

2. Helicopter — EXIT.

Rescue.

To open canopy and remove occupants:

1. Canopy door handles — ROTATE
(downward).

2. Lap belt/shoulder harness — RELEASE.

3. Helmet - DISCONNECT (HSS/ICS).

4. Occupants — REMOVE.

Change 1

5-5

V A

Section V
Part 1

NAVAIR 01-H1AAB-1

CANOPY JETTISON
DOOR ACTUATOR
DOOR HANDLE

DOOR SWITCHES
DOOR LOCK

MANUAL CLUTCH RELEASE
CLUTCH RESET

Removes glass from doors and windows.

Position door.

Latches door in closed position and deactivates electrical
circuit.

Opens or closes door electrically.

Secure helicopter doors.

Releases clutch for manual operation of door.

Sets clutch for electrical operation of door.

210900-118

Fig. 5-1. Emergency Egress and Rescue

NAVAIR 01 -HI AAB-1

Section V
Part 1 — Part 2

If canopy doors cannot be opened manually:

1. External canopy jettison handle access door
- PULL.

2. Handle ROTATE (90
counterclockwise) and PULL.

degrees

Personnel positioned within 50 feet of
helicopter could be injured by debris.

3. Occupants — REMOVE.

HOT START.

Indications (affected engine).

1. INLET TEMP exceeds 900 degrees C for 2
seconds, or

2. INLET TEMP exceeds 1150 degrees C.

Procedure.

1, Throttle — CLOSE.

2. FUEL - OFF.

3. START — ENERGIZE (for 30 seconds or

until INLET TEMP is below 300 degrees C).

4. Helicopter — SHUTDOWN.

5. Helicopter — EXIT.

ENGINE FIRE ON START (EXTERNAL).

Indications.

1. FIRE PULL warning light.

2. Smoke.

3. Fire.

Procedure.

1. Throttles — OFF.

f ■—

2. START - OFF.

3. FIRE PULL handle (affected engine) —

PUL T T --

4. FIRE EXT — MAIN/RESERVE.

—P——— ..... I .

5. Helicopter — SHUTDOWN.

6. Helicopter — EXIT.

PART 2 — TAKEOFF EMERGENCIES

SINGLE ENGINE FAILURE DURING
TAKEOFF.

Indications.

1. Left yaw.

2. RPM caution light.

3. MASTER CAUTION light.

4. Caution lights.

5. Rotor rpm decrease.

6. Engine instruments (affected engine)
decrease.

Procedure.

Gross weight, temperature, altitude, and airspeed
will determine if flight can be continued.

1, Collective — ADJUST (To maintain rpm and

desired power.)

2. Wing stores — JETTISON (as appropriate).
3 RPM - FULL INCREASE.

If insufficient altitude exists to continue flight:
1. Ground speed — DECREASE.

Change 1

5-7

Section V
Part 2 - Part 3

NAVAIR 01 -HIAAB-1

2. Landing attitude — ASSUME.

3. Collective — INCREASE (just prior to ground
contact to cushion landing).

4. Helicopter — SHUTDOWN.

If altitude permits, adjust airspeed for maximum
rate of climb or minimum rate of descent. After
gaining sufficient altitude or establishing
minimum rate of descent:

1. Affected engine — SECURE.

2. MASTER CAUTION light — RESET.

| 3. LAND AS SOON AS POSSIBLE.

DUAL ENGINE FAILURE DURING
TAKEOFF.

Indications.

1. Rapid settling.

2. Both engines instruments decrease.

3. Left yaw. *

'4. RPM caution light and audio.

5. MASTER CAUTION light.

6. ROTOR RPM decrease.

7. Caution lights.

Procedure.

When two-engine failure is experienced:

1. Autorotation — ESTABLISH.

NOTE

Gross weight, temperature, altitude, and
airspeed will determine if autorotation
can be established.

CAUTION

< > <;

Ground contact should be in a level
attitude to minimize helicopter damage.

The rotor brake should be applied to stop
the rotor prior to crew exiting the
helicopter.

PART 3 — INFLIGHT EMERGENCIES

HYDRAULIC MALFUNCTIONS.

Hydraulic System No. 1 Failure.

Indications.

1. Grinding or howling noise from pump.

2. Fluctuating or low hydraulic system
pressure.

3. MASTER CAUTION light.

4. No. 1 HYD PRESS caution light.

5. High tail rotor pedal force.

6. YAW SC AS disengaged.

. WHHVHHWU W W »

CAUTION

<; ' I

Cyclic and collective rate limiting
and/or control feedback may be evident
during abrupt maneuvers.

5-8

Change 1

NAVAIR 01-H1AAB-1

Procedure.

1. SCAS (YAW channel) - OFF.

2. HYD - SYS 1 OFF.

3. HYDR CONT circuit breaker — IN.

4. MASTER CAUTION light - RESET.

5. LAND AS SOON AS POSSIBLE.

Hydraulic System No. 2 Failure.

Indications.

1. Grinding or howling noise from pump.

2. Fluctuating or low hydraulic system
pressure.

3. MASTER CAUTION light.

4. NO. 2 HYD PRESS caution light.

5. PITCH and ROLL SCAS disengaged.

CAUTION

Cyclic and collective rate limiting
and/or control feedback may be evident
during abrupt maneuvers.

Procedure.

1. SCAS (PITCH and ROLL channels) -
OFF.

2. HYD - SYS 2 OFF.

3. HYD CONT circuit breaker - IN.

4. MASTER CAUTION light - RESET.

5. LAND AS SOON AS POSSIBLE.

Hydraulic Actuator/Servo Malfunctions.

The hydraulic system consists of two completely
independent power control subsystems. If an
actuator servo valve becomes inoperative, such as
foreign material causing a valve to jam, the
emergency servo valve bypass is actuated through

Section V
Part 3

pilot control inputs to maintain hydraulic powered
flight control. However, the pilot control force
required to accomplish the bypass operation in the
affected actuator will be higher than normal and
should cue the pilot that a hydraulic malfunction
has occurred. This increase in force will be noted
only in the control axis powered by the
malfunctioning actuator. Hydraulic system
pressure will remain normal, but a system
operating in the bypass mode may cause
overheating and overtemperature condition in the
affected system (Hydraulic System 1 or System 2).
This malfunction should be treated as an
individual system failure and the pilot should
follow the procedure for a single hydraulic system
failure.

If a HYD TEMP caution light is illuminated and
prolonged operation is necessary to reach a safe
landing area, the affected system should be turned
off to prevent further overheating; the system
could then be turned on again for the short period
of time for the landing procedure.

Indications.

1. Erratic control inputs.

2. Intermittent uncalled-for control inputs.

3. Abnormally high control force in a single
axis.

4. MASTER CAUTION light.

5. HYD TEMP caution light.

Procedure.

1. SCAS - OFF.

2. MASTER CAUTION light - RESET.

3. LAND AS SOON AS POSSIBLE (sliding
landing).

Complete (Dual) Loss of Flight Control
Hydraulic Boost (System 1 and System 2).

A safe recovery and landing from this type of
malfunction can be achieved provided the
following favorable conditions are satisfied:

NAVAIR 01 -HIAAB-1

VA Pai

I
|
I

Section V
Part 3

1. Helicopter attitude control is maintained.
Although flight control forces are
manageable by single pilot effort, the
transition from a power boosted to a non¬
power boosted flight control system could be
critical if encountered during high
performance maneuvers.

2. A suitable landing site is available;
preferably a hard surfaced runway (at least
3000 feet) with a long, shallow approach
capability.

Once stabilized helicopter attitude control is
achieved, abrupt control movements or maneuvers
should be avoided. Control movements will result
in normal flight reactions in all respects except for
the increased force required for the control move¬
ment. Flight control force characteristics are as
follows:

FORE AND AFT CYCLIC.

Nose down (forward
than nose up.

cyclic) stick force higher

WARNING

Pitch rates in excess of 3 degrees/
seconds should be avoided.

LATERAL CYCLIC.

Right roll force higher than left roll.

WARNING

Roll rates in excess of 3 degrees/seconds
should be avoided.

PEDALS.

Left pedal force slightly higher than right. If a yaw
oscillation develops, establish a steady state right
side slip attitude (one half ball width right).

COLLECTIVE.

Collective travel may be restricted to (approxi¬
mately) minimum of 30 percent torque and a
maximum of 50 percent torque and will become
increasingly difficult to move as each extreme of
displacement is approached.

The airspeed should be adjusted to 100-120 KIAS
for continued flight (return to base).

When landing without hydraulic boost, it is
recommended that very shallow approach to a
sliding landing be accomplished on a smooth,
hard surface. The approach should be initiated
from a straight-in position, 500 feet AGL or less,
and 2 nmi from touchdown point. Ideally, the
approach should be flown so .as to touchdown
at a minimum 20 KIAS with adequate margin
for the landing slide and stop; the primary flight
objectives will be to keep control movements to
a minimum but still maintain the airspeed and
sink rate that will terminate in a successful landing.

At airspeeds below 20 KIAS cyclic feed¬
back may be encountered. Do not
attempt to dampen feedback.

NOTE

At light gross weight configurations the
minimum power obtainable
(approximately 30 percent torque) may
not result in the desired sink rate unless
airspeed is reduced below 35 KIAS. In
this instance, a gross weight of 11,500 to
13,000 pounds would be desirable, so the
decision to retain wing stores should be
judiciously weighed in view of the
possibility of the requirement to wave off
the approach. Reduction of the throttles
may be utilized to establish a rate of
descent and minimize operation in the
shaded area of the Height Velocity
diagram.

Throttles should be increased to 100%
Nr prior to touchdown.

Indications.

1. MASTER CAUTION light.

2. HYD PRESS caution lights.

3. HYD PSI gages low.

4. All SCAS channels disengaged.

5. Increased control forces.

NAVAIR 01-H1AAB-1

Procedure.

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CAUTION

• :!

{ »»»»»»+»»»»»»»»»»»»» » »»

Avoid over-control and abrupt
movements.

1. Airspeed - ADJUST (to 100 KIAS for
continued flight).

2. SCAS (all channels) — CHECK OFF.

3. ENG RPM (N f ) — 100 PERCENT.

4. Wing stores — JETTISON (if necessary).

5. MASTER CAUTION light - RESET.

NOTE

Investigate collective limits.

6. Landing site — EVALUATE.

Ideally, a landing site with a hard surface and the
capability for a long shallow approach should be
selected. Initiate approach at approximately 2 nmi
from touch-down point and at an altitude of not
greater than 500 feet AGL.

7. Collective — DECREASE (to minimum

obtainable). *

8. Airspeed — ADJUST (to attain 300-500 fpm
rate of descent).

AHHWWVHHHVW ;

CAUTION

At very low gross weight, it may be
necessary to decrease airspeed to 35
KIAS or less to achieve 300-500 fpm rate
of descent. In the airspeed range of 25-35
KIAS, it will be necessary to decrease
airspeed to increase rate of descent. It
will also be necessary to increase
airspeed to decrease rate of descent.

NOTE

At high gross weights, the desired rate of
descent should be easily attained within
the obtainable power range.

Section V
Part 3

Attitude should be maintained once a 300 — 500
fpm rate of descent is achieved. As the landing
point is approached:

9. Rate of descent — MAINTAIN DESIRED
RATE OF DESCENT WITH
LONGITUDINAL CYCLIC.

Prior to touchdown:

10. ENG RPM (Nf) -100 percent.

11. Sliding landing - EXECUTE.

Rpm is necessary to maintain directional
control during the landing slide. Rolling
off throttles after touchdown will result
in decrease of directional control.

»W W< W >Wt » WW >W M

CAUTION

Since it will not be possible to move the
collective full down, the landing slide
will be very long. In zero wind conditions
it will be necessary to hold left cyclic
during the slide in order to maintain
lateral position.

12. Helicopter - SHUTDOWN. (Hold collective
at minimum until rotor stops.)

Wave-off With Complete Hydraulic Failure.

1. ENG RPM (Nf) -100 percent.

2. Power — INCREASE POWER (sufficiently to
clear obstacles and obtain a positive rate of
climb).

3. Airspeed - ADJUST TO 70 KNOTS
MINIMUM.

SCAS FAILURE

Indications.

1. Reduction in helicopter stability in affected
axis(es).

2. Increase in pilot workload to maintain
desired attitude.

3. Larger attitude deviations than desirable
with correction by the SCAS.

Change 1

5-11

ta p

Section V
Part 3

4. Erratic helicopter motion.

5. SCAS hardover will result in excessive roll,
pitch and yaw rates separately or together.

Procedure.

1. SCAS (Affected Channels) — CHECK OFF.

NOTE

If the helicopter pitches, rolls, or yaws
excessively without pilot input,
maintain control of helicopter and
disengage affected SCAS channel. If the
SCAS is not disengaged, the possibility
of the SCAS returning to the centered
position coupled with the pilot input to
stop the attitude excursion could result
in an overcontrolled helicopter response
in the opposite direction. When the
affected SCAS channel is disenaged, the
SCAS actuator will return to the
centered position almost
instantaneously; this coupled with a
simultaneous pilot input to stop the
attitude excursion could also result in an
overcontrolled helicopter response in the
opposite direction.

2. Airspeed — 100 KIAS.

3. LAND AS SOON AS PRACTICAL.

CONTROL SYSTEM MALFUNCTIONS.

Cyclic Control Interference.

Indications.

1. Stiffness or binding in control movement.

2. Restricted control travel.

Procedure.

1. Force trim - CHECK PROPER RELEASE.

NAVAIR 01-H1AAB-1

2. Control movements
MINIMUM.

- KEEP TO A

LAND AS
Landing).

SOON AS PRACTICAL (Sliding

Collective Control Interference.

Indications.

1. Stiffness or binding in control movement.

2. Restricted control travel.

Procedure.

1. Control movements — KEEP TO MINIMUM.

2. LAND AS SOON AS PRACTICAL (Sliding
Landing).

CAUTION

A shear pin is incorporated in the droop
compensator linkage connection to the
collective linkage. In case of a bind in the
droop compensator linkage, the shear
pin can be sheared to prevent binding of
the collective control. The droop
compensator is then inoperative and
extreme care must be taken to prevent
gas turbine overspeed and engine/rotor
underspeed.

TAIL ROTOR MALFUNCTION.

There is no single emergency procedure for all
types of anti-torque malfunctions. The key to a
pilots successful handling of a tail rotor
emergency lies is his ability to quickly recognize
the type malfunction that has occurred.

Complete Loss of Tail Rotor Thrust.

This is a situation involving a break in the drive
system, such as a severed driveshaft, wherein the
tail rotor stops turning and no thrust is delivered
by the tail rotor. A failure of this type in powered
flight will always result in the nose of the
helicopter swinging to the right (left sideslip) and
usually a roll of the fuselage. Nose-down tucking
will also be present. The most advisable procedure
is to reduce power, to engine idle if necessary, and
coordinate the resulting maneuver with cyclic
control. At some gross weights it is possible that a
stabilized powered flight condition can be
achieved if the loss of the tail rotor thrust occurs at
a high enough airspeed. Once stabilized in an
autorotation, some power may be applied (altitude
permitting) to see if powered flight is possible.

NAVAIR 01 -HI AAB-1

Section V
Part 3

For most gross weights, it is unlikely
that the AH-IT can achieve a stabilized
power flight condition following loss of
tail rotor thrust. Emphasis should be
placed on entering autorotation
immediately by reducing collective and
throttle setting. Control of heading will
probably not be regained in autorotative
flight. The pilot should expect that some
rotation will be present until touchdown.
Touchdown should be executed in as
level an attitude as can be achieved.
Ground speed should be as slow as
possible to minimize the possibility of
turnover.

Loss of Tail Rotor Components.

The loss of any tail rotor components will result in
a forward CG shift. Other than additional nose-
down tuck, this situation would be quite similar to
complete loss of tail rotor thrust as discussed
above.

Fixed Pitch Failures.

Failures of this type (broken control tubes, jammed
slider, etc.) are characterized by either a lack of
directional response when a pedal is pushed or the
pedals in a locked position. If the pedals cannot be
moved with a moderate amount of force, do not
attempt to apply a maximum effort since a more
serious malfunction could result. If the helicopter
is in a trimmed condition when the malfunction is
discovered, the engine power and airspeed should
be noted and the helicopter flown to a suitable
landing area. Combinations of engine torque,
rotor rpm, and airspeed will correct or aggravate a
yaw attitude and these are what will be used to
land the helicopter.

LEFT PEDAL APPLIED.

If the tail rotor pitch becomes fixed during a high-
power condition (left pedal applied), the helicopter
will yaw to the left when power is reduced. Under
these conditions, the power should be reapplied
and airspeed adjusted to a value where a
comfortable yaw angle can be maintained. If
airspeed is increased, the vertical fin will become
more effective and an increased left yaw attitude
will develop. To accomplish landing, establish a

power approach with sufficiently low airspeed
(zero, if necessary) to attain a rate of descent with a
comfortable side slip angle. As collective is
increased just before touchdown, left yaw will be
reduced.

NOTE

Use throttle, not RPM switch, for rpm
control.

RIGHT PEDAL APPLIED.

If the tail rotor pitch becomes fixed during cruise
flight or a reduced power situation occurs (right
pedal applied) the helicopter will yaw to the right
when power is increased. For either of these
situations, a sliding landing can be performed.
Throttles may be reduced as required when adding
collective at touchdown and cushion the landing
with collective. If the right yaw becomes excessive,
roll on the throttles and initiate a wave off. The
greatest problem is the compromise that may have
to be made between rate of descent and yaw
attitude since the collective (power) is the primary
control for both of these parameters. Within
reasonable limits, it is probably preferable to land
hard with a zero yaw attitude than to make a soft
landing while in a severe yaw attitude.

Emergency Procedure For
Malfunction While At A Hover.

Anti-Torque

1. In the event of complete loss of tail rotor
thrust or loss of tail rotor components, close
throttle and perform hovering autrotation.

2. In the event of loss of tail rotor pitch
control, close throttle and perform hovering
autorotation.

3. In the event of jammed tail rotor pitch
control, gradually reduce collective pitch to
accomplish a power touchdown.

MAST BUMPING

Indications:

1. Sharp two/rev knocking.

Procedure:

During high speed sideward and rearward flight:

1. Cyclic - IMMEDIATELY APPLY SLIGHTLY
TOWARD CENTER.

Change 1

NAVAIR 01 -HI AAB-1

Section V
Part 3

2. Pedals - IMMEDIATELY APPLY AS
REQUIRED TO BRING NOSE INTO RELA¬
TIVE WIND.

3. LAND AS SOON AS POSSIBLE.
During all other flight conditions:

1. Cyclic - IMMEDIATELY CENTER LAT¬
ERALLY, THEN AFT AS REQUIRED TO
MAINTAIN POSITIVE G LOAD ON ROTOR.

2. Controls - AS REQUIRED TO REGAIN
BALANCED FLIGHT.

3. LAND AS SOON AS POSSIBLE.

UNCOMMANDED RIGHT ROLL DURING
FLIGHT BELOW 1 G.

Indications:

1. Uncommanded right roll.

2. Reduced cyclic effectiveness.

Procedure:

1. Cyclic - IMMEDIATELY CENTER LATER¬
ALLY, THEN AFT.

Lateral cyclic* is decreasingly effective
below 1 G and increases main rotor
flapping which can result in mast
bumping. Do not engage SCAS during
recovery.

When main rotor returns to a positive thrust
condition:

2. Controls - AS REQUIRED TO REGAIN
BALANCED FLIGHT.

If mast bumping occurred:

3. LAND AS SOON AS POSSIBLE.

ENGINE MALFUNCTIONS.
Single Engine Failure (In Flight).

The pilot’s reaction to the failure of a single engine
encompasses two general areas; control of the

helicopter and possible engine restart. In all cases,
control of the helicopter, attitude, altitude and
rotor rpm should take precedence over any attempt
to restart a failed engine. Under high gross weight
and density altitude conditions, level flight may
not be possible. At maximum single engine power
available, and at low AGL altitude, the external
wing stores should be jettisoned to reduce gross
weight so that level flight can be achieved. This
should give the pilot sufficient time to analyze
possible causes of the failure and make a decision
whether or not to attempt an airstart. When one
engine fails, rotor speed can be expected to droop.
The desired rotor rpm can be regained if sufficient
power is available, by using the engine RPM
switch. After rpm is regained by use of the RPM
switch, desired rotor rpm can be maintained by the
collective control.

Indications.

1. Left yaw.

2. RPM caution light (gas producer).

3. MASTER CAUTION light.

4. Rotor rpm decrease.

5. Engine instruments decrease.

6. CAUTION panel lights.

Procedure.

1. Collective - ADJUST TO MINIMUM

REQUIRED.

2. Wing stores — JETTISON (if appropriate).

Under conditions of high gross weight or low
altitude and low airspeed, strong
consideration should be given to jettisoning
of wing stores simultaneously with step 1.

3. RPM — FULL INCREASE.

4. ECU and RAIN RMV — OFF.

5. Failed engine — IDENTIFY.

6. Throttle failed engine — CLOSE.

7. FUEL failed engine — OFF.

8. GEN failed engine — OFF.

9. PART SEP (normal engine)— OFF.

5-14

Change 1

Part 3

10. MASTER CAUTION light - RESET.

11. If desired — AIRSTART.

12. LAND AS SOON AS POSSIBLE.
Airstart.

- - 1

CAUTION

!> ; |

If the cause of failure is obviously
mechanical as evidenced by abnormal,
metallic, or grinding sounds, do not
attempt an airstart.

1. Throttle (affected engine) — CLOSED.

2. GEN-OFF.

3. GOV - AS DESIRED.

NOTE

The decision concerning which portion
of the fuel control to use must be made by
the pilot based on his analysis of the
engine failure and his own skill in flying
with the manual fuel control.

4. FUEL - ON.

5. FUEL PRESS - NORMAL.

6. ENGINE OIL TEMP-INDICATES NORMAL
OR LESS.

NOTE

• Energizing the starter generator on the
failed engine will result in an increased
power demand on the operating engine.
Turning the operating starter generator
off will alleviate the increase in power
demand.

• Abnormal instrument readings of the
failed engine may indicate that an
airstart might be inadvisable. INLET
TEMP of operating engine will rise
when attempting airstart.

7. START — ON (affected engine).

8. Engine oil PRESS - POSITIVE INDICATION.

9. Throttle - (With INLET TEMP below 200°C
and GAS PROD (Ng) at 12 percent or above)
OPEN (engine idle).

10. INLET TEMP - MONITOR.

11. START - OFF (at 50 percent Ng).

12. ENGINE OIL - CHECK PRESSURE AND
TEMPERATURE.

13. GEN - ON.

14. Throttle — INCREASE (to match engine
torques).

15. LAND AS SOON AS PRACTICAL.

Engine Shutdown In Flight.

If an airstart is unsuccessful or not desired:

1. Throttle - CLOSE.

2. FUEL - OFF.

3. START - OFF.

4. GEN - OFF.

5. LAND AS SOON AS PRACTICAL.

Single Engine Failure (Hovering In-Ground
Effect).

Indications.

Same as single engine failure in flight.

Procedure.

1. Heading and landing attitude —
MAINTAIN.

Single Engine Failure (Hovering Out-Of-
Ground Effect).

Indications.

Same as single engine failure in flight.

2. Collective — ADJUST (to control rate of
descent and cushion landing).

9 .

5-15 £

7 >

Change 1

Section V
Part 3

NAVAIR 01 -HI AAB-1

Procedure.

1. Heading and attitude control — MAINTAIN.

2. Collective — ADJUST (to maintain rpm and
desired power).

If insufficient power exists to fly away:

1. Attitude - ASSUME LAND ATTITUDE.

2. Collective — INCREASE (just prior to ground
contact to cushion landing).

If altitude permits, adjust airspeed for
maximum rate of climb or minimum rate of
descent. After gaining sufficient altitude or
establishing minimum rate of descent.

1. Affected engine — SHUTDOWN.

2. LAND AS SOON AS POSSIBLE.

Dual Engine Failure.

Under operational conditions, the altitude-
airspeed combination for a safe autorotative
landing is dependent upon many variables such as
pilot capabilities, density altitude, helicopter gross
weight, proximity of a suitable landing area, and
wind direction and velocity in relation to flight
path. This does not preclude operation in the
shaded area of the height velocity diagram under
emergency or pressing operational requirements.
Immediately upon a two-engine failure, rotor rpm
will decay and the nose of the helicopter will swing
to the left. This is due to the loss in power and
corresponding reduction in torque. Except in those
instances when a two-engine failure is
encountered in close proximity to the surface, it is
mandatory that autorotation be established by
immediately lowering the collective pitch to
minimum.

Heading can be maintained by depressing the
right pedal to decrease the tail rotor thrust.
Autorotative rpm will vary with different ambient
temperature, pressure altitude, increase in G
loading, and gross weight conditions. High gross
weights, increased G loads, and higher altitudes
and temperature will cause increased rpm which
can be controlled by increasing collective pitch.
Any increase of rotor rpm, other than specified for
maximum glide, will result in a greater rate-of-
descent. Therefore, if time permits, adjusting the

collective pitch lever to produce the desired rotor
rpm will result in an extended glide. At an altitude
of approximately 75-100 feet, a flare should be
established by moving the cyclic stick aft with no
change in collective pitch. This will decrease both
airspeed and rate-of-descent and cause an increase
in rotor rpm. The amount that the rotor rpm will
increase is dependent upon gross weight and the
rate that the flare is executed. An increase is
desirable because more energy will be available to
the main rotor when collective pitch is applied.

DUAL ENGINE FAILURE IN FLIGHT.

Indications.

1. Rapid settling.

2. Both engines’ instruments decrease.

3. Rotor rpm decreases.

4. Left yaw.

5. RPM caution light and audio.

6. MASTER CAUTION light.

7. CAUTION panel lights.

DUAL ENGINE FAILURE AT HIGH POWER
AND HIGH AIRSPEED.

1. Cyclic — IMMEDIATELY AFT.

NOTE

The aft cyclic input will command a nose
up attitude change, will initiate a flare to
reduce airspeed and will maintain a
rotor loading. If the aft cyclic command
is not made, the aerodynamic loading
will result in an uncommanded nose up
pitch change. The SCAS will detect this
as any other external disturbance, and
will compensate with a nose down pitch
correction.

2. Collective — DECREASE.

CAUTION

A SCAS nose-down pitch correction
combining with rapid decrease of col¬
lective could cause less than +0.5 G load¬
ing resulting in excessive main rotor
flapping and possible mast bumping.

5-16 Change 1

3. Autorotation at (65-120 KIAS)

ESTABLISH.

DUAL ENGINE
ALTITUDE.

ZERO TO 50 KIAS, 20 FEET ALTITUDE OR
BELOW. From this condition of airspeed and low
altitude, flare capability is limited and caution
should be exercised to avoid striking the ground
with the tail; the primary objective is to level the
skids prior to ground contact. Initial collective
reduction varies with altitude; from a 4-foot skid
height, do not attempt collective reduction but use
the available rotor energy and collective to
cushion touchdown; above 4-foot skid heights, a
partial reduction of collective will maintain rotor
rpm until up collective is initiated to cushion
touchdown.

50 TO 70 KIAS, 20 FEET ALTITUDE AND
BELOW. From this condition flare capability is
good. Initiate a cyclic flare and reduce collective to
maintain rotor rpm, minimize rate of descent, and
decelerate helicopter; level skids prior to ground
contact and utilize collective to cushion
touchdown.

NOTE

The optimum flare airspeed for all gross
weights is 75 KIAS.

75 KIAS TO Vh AIRSPEED, 20 FEET
ALTITUDE AND BELOW. Immediately execute a
cyclic flare to initiate a climb to 25 feet or higher
and lower collective as necessary to maintain rotor
rpm; achieve 75 KIAS and maintain until normal
flare and touchdown is accomplished.

Procedure.

Autorotation - ENTER IMMEDIATELY.

If conditions permit:

DUAL ENGINE FAILURE (HOVERING IN-
GROUND EFFECT).

FAILURE AT LOW , ndications .

2. Airstart — ATTEMPT
engines).

(on one or both

If airstart is successful, follow procedures for
single engine landing.

If airstart is unsuccessful, follow procedures for
autorotative landing.

Same as indications for Dual Engine Failure in
Flight.

Procedure.

1. Directional
HEADING.

pedals — MAINTAIN

2. Attitude — LEVEL (to

pitching on touchdown).

prevent adverse

3. Collective — INCREASE (to cushion

landing).

Upon touchdown:

4. Cyclic - CENTERED.

5. Collective — FULL DECREASE.

6. Rotor brake handle — ENGAGE.

CAUTION

Regardless of sink rate at touch-down,
damage will be minimized when in a
level attitude.

DUAL ENGINE FAILURE (HOVERING OUT-
OF-GROUND EFFECT).

Indications.

Same as indications for Dual Engine Failure in
Flight.

Procedures.

1. Collective — DECREASE (to maintain rotor

rpm).

D irectional

HEADING.

pedals — MAINTAIN

3. Attitude — LOWER NOSE TO INCREASE

AIRSPEED (if possible).

NOTE

If altitude permits attempt to attain
optimum autorotation flare airspeed.

Prior to touchdown.

4. Collective — INCREASE (to cushion
landing).

5. Cyclic — AS REQUIRED (to level helicopter).

CAUTION

Regardless of sink rate at touchdown,
damage will be minimized when in a
level attitude.

Upon touchdown:

6. Collective — FULL DECREASE.

7. Rotor brake handle - ENGAGE.

Power Turbine Governor (NF) Failure.

Indications.

1. Erratic GAS PROD RPM (Ng).

2. Erratic INLET TEMP.

3. Fluctuating ENG RPM (Nf).

4. Abrupt increase in ENG RPM (Nf) above
governed value.

5. Abrupt decrease in ENG RPM (Nf) below
governed value.

6. Fluctuating TORQUE.

Procedure.

1. Affected engine — IDENTIFY.

2. Throttle — ENGINE IDLE.

3. GOV — MANUAL. (It is not necessary to
wait for GAS PROD (Ng) to stabilize at
engine idle before switching to manual.)

CAUTION

The engine being operated on manual
fuel control shall be closely monitored to
ensure limits are not exceeded.

4. Throttle — ADVANCE (slowly to desired
power setting).

5. LAND AS SOON AS PRACTICAL. •

Engine Underspeed Gas Prod (Ng).

Indications.

1. Abrupt decrease in GAS PROD (Ng).

2. Subsequent decrease in ENG RPM (Nf).

3. Possible decrease in ROTOR RPM (Nr).

4. Decrease in TORQUE (affected engine).

Procedure.

1. Collective — ADJUST (maintain 97 - 100%

ROtOR RPM (Nr)). '

2. Affected engine — IDENTIFY.

3. Throttle - ENGINE IDLE.

4. GOV - MANUAL.

CAUTION

The engine being operated on MANUAL
fuel control shall be closely monitored to
ensure that limits are not exceeded.

5. Throttle — ADVANCE (slowly to desired
power setting).

6. LAND AS SOON AS PRACTICAL.

Engine Overspeed Rotor RPM (NR).
Indications.

1. Increase in ROTOR RPM.

2. Increase in ENG RPM (Nf) (affected engine).

3. Increase in GAS PROD RPM (Ng) (affected
engine).

4. Increase in TORQUE (affected engine).

5-18

NAVAIR 01 -HI AAB-1

CAUTION

The increase in the above parameters of
the affected engine may result in a
corresponding decrease in these
parameters on the normal engine. A
cross check of both engines’ gages will
preclude a false diagnosis of a power loss
of the engine not experiencing the
overspeed.

Procedure.

1. Collective — ADJUST (to avoid overspeed).

2. Affected engine — IDENTIFY.

3. Throttle - ENGINE IDLE.

4. GOV - MANUAL.

CAUTION

1 1

The engine being operated on MANUAL
fuel control shall be closely monitored to
ensure that limits are not exceeded.

If unable to control engine, secure engine,
control is attained:

5. Throttle — ADVANCE (slowly to desired
power setting).

6. LAND AS SOON AS PRACTICAL.

Engine Fire in Flight.

Indications.

1. Smoke.

2. Fumes.

3. Fire.

4. FIRE PULL warning lights.

Procedure.

NOTE

Fire may be confirmed by yawing the
helicopter and observing the smoke
trail.

Section V
Part 3

1. Throttle - CLOSE.

2. FIRE PULL handle - PULL.

3. FIRE EXT — MAIN (if fire indications

persists, switch extinguisher to RESERVE.) .

4. Single engine procedure — EXECUTE.

5. CROSSFEED - CLOSE.

6. FUEL BOOST circuit breaker — PULL.

NOTE

FUEL AFT BOOST circuit breaker shall
be pulled for fire in No. 2 engine and
FUEL FWD BOOST circuit breaker
shall be pulled for fire in No. 1 engine.
With cross feed valve closed and FUEL
AFT BOOST circuit breaker pulled,
there will be no indication of fuel
pressure.

7. MAYDAY - BROADCAST.

8. LAND AS SOON AS POSSIBLE.

9. If fire persists — LAND IMMEDIATELY.

Fire-Both Engines In Flight.

In the event that both FIRE PULL warning lights

Indications.

1. Smoke.

2. Fumes.

illuminate simultaneously in flight, a decision
must be made whether or not to terminate the
subsequent approach with a full autorotational
landing or with a power recovery and landing.
This decision will be based on the length of time
required to land the helicopter; i.e., the extent to
which the fire will spread prior to landing. At
higher altitudes, it may be necessary to secure both
engines in order to extinguish the fire before
incurring catastrophic damage. At lower altitudes,
it may be more prudent to land with power (on at
least one engine) in order to increase the
probability of a safe landing. In any event, an
immediate landing must be made, and the
ultimate decision on how far to proceed beyond
Step 1 (below) must be based on altitude and rests
with the pilot in command.

Section V
Part 3

NAVAIR 01 -HI AAB-1

3. Fire.

4. FIRE PULL warning lights.

Procedure.

1. Autorotation — ENTER.

2. Throttles - CLOSE.

3. FIRE 1 PULL handle - PULL.

4. FIRE EXT - MAIN.

5. FUEL (both engines) - OFF.

6. FIRE 1 PULL handle - PUSH IN .

7. FIRE 2 PULL handle - PULL.

8. FIRE EXT - RESERVE.

9. MAYDAY — BROADCAST.

10. Autorotative landing — ACCOMPLISH.
Engine Chip Detr Caution Light.

Indications.

1. CHIP DETR caution light (affected engine).

2. MASTER CAUTION light.

Procedure.

In the event of an engine CHIP DETR light with
no secondary indications, consideration should be
given to applicable single engine procedures
without securing the affected engine. The affected
engine may be advanced to accomplish a safe
landing.

1. Throttle - ENGINE IDLE.

2. LAND AS SOON AS PRACTICAL.

If secondary indications exist:

1. Single engine procedure — EXECUTE.

Engine Oil Pressure Low.

Indications.

1. ENGINE OIL PRESS decreases (affected
engine).

2. OIL PRESS caution light (affected engine).

3. MASTER CAUTION light.

Procedure.

1. ENGINE OIL PRESS - CHECK.

If oil pressure is above lower limit:

1. MASTER CAUTION light - RESET.

2. ENGINE OIL P gage - MONITOR.

3. LAND AS SOON AS PRACTICAL.

If oil pressure is below lower limit:

1. Throttle — CLOSE.

2. Single engine procedures — EXECUTE.

Engine Oil Overtemperature.
Indications.

1. ENGINE OIL TEMP high.

Procedure.

1. ENGINE OIL PRESS - CHECK.

If oil temperature is above 116 and oil presssure
is below 40 psi:

1. Engine - SHUTDOWN.

2. Single engine procedure — EXECUTE.

3. LAND AS SOON AS POSSIBLE.

Compressor Stalls.

Indications.

1. No throttle response.

2. High or erratic INLET TEMP.

3. Decreasing or erratic GAS PROD (Ng) and
ENG RPM (Nf).

4. Rapid engine “chugs” or explosions.

5-20

NAVAIR 01 -HIAAB-1

Section V
Part 3

Procedure.

1. Throttle — DECREASE.

2. ! INLET TEMP - MONITOR.

If compressor stall persists:

1. Throttle - CLOSE.

2. Single engine procedure — EXECUTE.

3. AMPS indication of zero.

Procedure.

1. GEN BUS RESET circuit breaker - CLOSED.

2. GEN FIELD circuit breaker — CLOSED.

3. GEN — MOMENTARY RESET, THEN ON.

4. MASTER ARM - OFF.

5. MASTER CAUTION light - RESET.

MAIN DRIVESHAFT FAILURE.

A main driveshaft failure presents the pilot with
confusing aural and visual cues that require
prompt interpretation and corrective action. ENG
RPM (Nf) will indicate over speeding, but ROTOR
RPM (Nr) will decay rapidly and the helicopter will
yaw left; the low rotor RPM caution light and
audio signal will be activated. Immediate response
to the low rotor RPM caution is required to prevent
an excessively low rpm situation from which a safe
recovery would be extremely difficult.

Indications.

1. High ENG RPM (Nf).

2. Low rotor RPM caution light and audio.

3. Zero TORQUE indication.

4. Grinding sounds.

If the mission does not require the use of the
MASTER ARM switch, complete the mission.

Failure of One Generator (MASTER ARM
Switch Required).

Indications.

1. MASTER CAUTION light.

2. DC GEN caution light.

3. AMPS indication of zero.

Procedure.

1. GEN BUS RESET circuit breaker — CLOSED.

2. GEN FIELD circuit breaker - CLOSED.

3. GEN — MOMENTARY RESET, THEN ON.

4. MASTER ARM - OFF.

Procedure.

NOTE

1. Autorotation — ENTER IMMEDIATELY.

2. Throttles - FLIGHT IDLE.

If time, altitude, and conditions permit:

3. Throttles — CLOSE.

4. FUEL switches — OFF.

ELECTRICAL SYSTEM MALFUNC¬
TIONS.

Failure of One Generator (MASTER ARM
Switch Not Required).

Indications.

1. MASTER CAUTION light.

2. DC GEN caution light.

The MASTER ARM switch should be
kept OFF as much as possible so that the
remaining generator can keep the
battery charged. With MASTER ARM
switch in STBY or ARM, the remaining
generator furnishes armament power
and the battery furnishes helicopter’s
power. As battery voltage is depleted,
the SC AS performance will be degraded
and may disengage.

5. NON-ESS BUS - MANUAL.

NOTE

With the non-essential bus switch in the
NORMAL position, the main inverter,
the TACAN and the air-conditioner will
not be operable. The standby inverter
will automatically come on the line and
will be powered by the battery.

Change 1

NAVAIR 01-H1AAB-1

Section V
Part 3

6. TACAN - AS REQUIRED.

7. ECU/VENT - OFF.

8. Unnecessary equipment — OFF.

9. MASTER CAUTION light - RESET.

10. LAND AS SOON AS PRACTICAL.

Failure of Both Generators.

In the event both generators fail in flight,
emergency dc power is supplied by a 24 volt, 34
ampere hour battery. This battery, assuming an 85
percent charge, can supply the essential bus for a
period of approximately 32 minutes. With the
NON-ESS BUS in MANUAL position, the battery
will supply the non-essential bus under emergency
conditions for a period of approximately 16
minutes. To conserve battery power, all unneeded
navigation equipment and radios should be
secured.

Indications.

1. AMPS indicate zero.

2. DC GEN caution light.

3. MASTER CAUTION light.

Procedure.

1. GEN — OFF.

2. NON-ESS BUS — MANUAL.

3. GEN BUS RESET circuit breaker - CLOSED.

4. GEN FIELD circuit breaker - CLOSED.

5. All unnecessary equipment — Off.

6. GEN - RESET, THEN ON.

7. MASTER CAUTION light - RESET.

8. LAND AS SOON AS PRACTICAL.

If generator does not come on:

1. GEN — OFF.

2. LAND AS SOON AS PRACTICAL.

Main Inverter Failure.

Indications.

1. MASTER CAUTION light.

2. AC MAIN caution light.

3. Non essential ac bus functions cease.

4. INV MAIN circuit breaker — OPEN.

NOTE

Loss of dc power from the non-essential
bus to the main inverter will result in
automatic switch over to the standby
inverter. If the main inverter fails
internally while dc power is applied, the
automatic switch over will not occur.

Procedures.

1. INVERTERS - STBY.

2. INV MAIN circuit breaker — OPEN.

3. MASTER CAUTION light - RESET.

4. SCAS - ENGAGED.

5. LAND AS SOON AS PRACTICAL.

Failure of Both Inverters.

Indications.

1. MASTER CAUTION light.

2. AC STBY caution light.

3. All ac bus functions cease.

4. INV STBY circuit breaker — OPEN.

Procedure.

1. INV STBY circuit breaker — CLOSED.

2. INV MAIN circuit breakers — CLOSED.

3. INVERTERS — MAIN.

5-22

NAVAIR 01 -HIAAB-1

If either inverter functions:

1. Unnecessary equipment — OFF.

2. MASTER CAUTION light - RESET.

3. LAND AS SOON AS PRACTICAL.

If neither inverter functions:

1. INVERTERS — OFF.

2. Unnecessary equipment — OFF.

3. MASTER CAUTION light - RESET.

4. LAND AS SOON AS PRACTICAL.
Complete Electrical Failure.

CAUTION

If one or both engines are in manual fuel
and power is removed from the 28 vdc
essential bus, the fuel control solenoid
will de-energize and actuate the
automatic fuel control regardless of
governor switch position.

Indications.

1. All electrical functions cease.

CAUTION

' iwwwwwwwmwM >

Total loss of electrical power will cause
the loss of all engine and component
instruments, indicators, gages and tacho¬
meters except for GAS PROD (Ng)
tachometers.

Procedure.

1. Copilot/gunner ELEC PWR — ELEC PWR.

2. Airspeed — REDUCE (100 KIAS or less).

3. LAND AS SOON AS PRACTICAL.

Section V
Part 3

Battery Overtemp/Thermal Runaway
Indications.

1. MASTER CAUTION light.

2. BATTERY TEMP light.

3. High AMPS indication.

4. Smoke or fumes emitting from battery
compartment.

5. Muffled bang or thud sound in battery
compartment.

Procedure.

1. BATTERY - OFF.
If on deck:

1. Helicopter — SHUTDOWN (alert crash
crew).

If in flight:

1. LAND AS SOON AS POSSIBLE.

2. Helicopter
crew).

SHUTDOWN (alert crash

WARNING

Do not use fire extinguisher on battery
if there is no visible fire because it may
cause an explosion. If a visible fire has
developed, fire extinguisher may be used.

Indications.

1. Smoke or fumes in cockpit.

2. Equipment failure.

Procedure.

ELIMINATION OF SMOKE AND FUMES
IN COCKPIT.

-23 f

1. ECU/VENT - OFF.

Section V
Part 3

NAVAIR 01-H1AAB-1

Canopy doors — OPEN
INTERMEDIATE POSITION (at a
maximum of 45 KIAS, to eliminate excessive
smoke or fumes).

Do not discharge hand portable fire
extinguishers in closed cockpit due to the
possibility of disabling the crew.

ELECTRICAL FIRE.

Do not discharge hand portable fire
extinguishers in closed cockpit due to the
possibility of disabling the crew.

Indications.

1. Smoke or fumes.

2. Equipment failure.

3. Fire.

4. High AMPS indication.

Procedure.

1. Both GEN — OFF.

CAUTION

4 !wh% hh%hh h h h»

Do not attempt target run with less than
one generator and battery.

NOTE

With both Generators off and the NON-
ESS BUS switch in NORMAL, the
inverter function automatically
switches to the standby inverter.

2. Circuit breakers — CHECK.

CAUTION

,» <;

Do not reset any circuit breakers that are
tripped. It is likely that those circuits are
the problem.

3. All unnecessary equipment — OFF.

4. MASTER CAUTION light - RESET.

If fire is not evident and/or an immediate landing
is not possible:

5. No. 1 GEN — ON.

6. No. 2 GEN - ON.

7. INVERTERS - MAIN.

8. Necessary equipment — ON.

If fire is evident on any step 5 through 8:

9. Applicable equipment — OFF.

10. Applicable circuit breaker — OPEN.

11. LAND AS SOON AS PRACTICAL.

If evidence of fire persists:

12. Both GEN — OFF.

13. BATTERY — ON (only as required).

14. LAND AS SOON AS POSSIBLE.

CAUTION

SCAS will disengage with no electrical
power.

FUSELAGE FIRE IN FLIGHT.

Procedure.

Do not discharge hand portable fire
extinguishers in closed cockpit due to the
possibility of disabling the crew.

1. LAND IMMEDIATELY.

Change 1

NAVAIR 01-H1AAB-1

Section V
Part 3

FUEL SYSTEM MALFUNCTIONS.

Fuel Boost Pump Failure.

The helicopter is equipped with two electrically
driven fuel boost pumps, either of which is capable
of supplying sufficient fuel to both engines. A
complete helicopter fuel system failure will not be
common because of separate forward and aft fuel
boost pumps.

Avoid helicopter operation with dual
fuel boost pump failure above 6000 ft.
pressure altitude. This can result in an
engine flame-out due to fuel starvation.

CAUTION

i ►

With forward boost pump inoperative, a
nose down attitude in excess of 14° will
result in 538 pounds or more of
unuseable fuel. Flame-out could result.

Indications.

1. FWD or AFT FUEL BOOST caution light.

2. MASTER CAUTION light,

3. A decrease in fuel pressure.

4. FUEL BOOST circuit breaker — OPEN.

Procedure.

1. TANK INTCON — OPEN.

2. CROSS FEED - OPEN.

NOTE

With cross feed closed, fuel pressure
gage indicates aft fuel boost pressure
only.

3. FUEL BOOST circuit breaker — OPEN.

4. MASTER CAUTION light - RESET.

5. LAND AS SOON AS PRACTICAL.

Engine Driven Fuel Pump Failure.

If the engine driven fuel pump fails, the engine will
flame-out due to fuel starvation.

Indications.

1. Engine instruments decrease.

2. Rotor rpm decreases.

3. MASTER CAUTION light.

4. Caution panel lights.

Procedure.

1. Follow procedure for single engine failure in
this section.

Engine Fuel Filter Impending Bypass.
Indications

1. FUEL FILTER caution light (affected
engine).

2. MASTER CAUTION light.

Procedure.

1. MASTER CAUTION light — RESET.

2. LAND AS SOON AS PRACTICAL.

IMPENDING TRANSMISSION FAILURES.

An impending transmission failure may be
indicated by any unusual noise or vibrations from
the transmission area, abnormal transmission oil
pressure or temperature indications, transmission
oil bypass light, transmission chip light, loss of
Nr, or yaw kicks. These indications may occur

Change 1

5-25

IMAVAIR 01 -HIAAB-1

Section V
Part 3

singularly or in combination. Generally, there are
two extremes that can be expected in a trans¬
mission failure; seizure of the drive train or a
disconnect of the drive train that would allow the
rotor system to turn independently of the engines.
If an impending transmission failure is suspected,
whether it is due to oil starvation or a power
discontinuity, priority must be given to maintaining
Nr, descending, and landing as soon as possible.
Nr may be maintained by using a combination of
collective pitch setting (with throttles full open)
and airspeed. A smooth transition to an airspeed
providing minimum power requirements should be
| accomplished. Aircraft controllability will become
markedly degraded if Nr decreases below 90
percent.

Indications.

1. Unusual noise or vibrations from transmission
areas.

2.XMSN OIL HOT caution
OIL TEMP gage high.

light with XMSN

3. XMSN OIL PRESS caution light with XMSN
OIL PRESS gage low.

4. XMSN OIL BYP caution light with XMSN oil
temp gage high.

5. XMSN CHIP DET caution light.

6. MASTER CAUTION LIGHT.

7. Nr low.

8. Yaw kicks.

Procedures.

1. Maintain Nr.

2. Descend — POWER ON.

3. LAND AS SOON AS POSSIBLE.

| If Nr decays or more violent vibrations occur:

4. Lower collective with throttles full open.

5. Maintain powered descent.

6. LAND IMMEDIATELY.

If continued flight is mandatory and appears
feasible:

1. Establish slow flight of 50 KIAS and 20 feet
AGL.

If Nr decays or more violent vibrations occur:

2. Maintain Nr.

3. LAND IMMEDIATELY.

• If an engine to transmission disconnect
occurs, Nf may tend to overspeed.
Priority must be given to maintaining
Nr before attempting to control the Nf
overspeed.

• Autorotation in the event of transmission
oil starvation may contribute to trans¬
mission seizure.

• In certain modes of transmission failures,
loss of hydraulic systems or tail rotor
drive may occur.

CAUTION

( ' <|

• With indications of an impending trans¬
mission failure, an approach should be
made with minimum power changes to
minimize the chance of seizure. Control
movements should also be kept to a
minimum.

• Because of the “wet bulb” temperature
system, oil starvation may not be
accompanied by a rising temperature.

COMBINING GEARBOX MALFUNC¬
TIONS.

Power shall be maintained throughout
approach and landing to aid in
preventing seizure of gears.

NAVAIR 01 -HIAAB-1

Combining Gearbox Oil Overtemperature
Indications.

1. C BOX OIL HOT caution light.

2. MASTER CAUTION light.

3. GEARBOX OIL TEMP high.

Procedures.

Section V
Part 3

1. Collective — DECREASE.

2. GEARBOX OIL TEMP - CHECK.

3. GEARBOX OIL PRESS - CHECK.

4. MASTER CAUTION light - RESET.

5. LAND AS SOON AS PRACTICAL.

If GEARBOX OIL TEMP indication is above
limit, or if GEARBOX OIL PRESS is below
limit:

1. LAND AS SOON AS POSSIBLE.

Combining Gearbox Oil Pressure Low.
Indications.

1. C BOX OIL PRESS caution light.

2. MASTER CAUTION light.

3. GEARBOX OIL PRESS indicates low.
Procedure.

1. Collective — DECREASE.

2. GEARBOX OIL PRESS — CHECK.

3. GEARBOX OIL TEMP — CHECK.

4. MASTER CAUTION light - RESET.

5. LAND AS SOON AS PRACTICAL.

Combining Gearbox Chip Detector.
Indications.

1. C BOX CHIP DETR caution light.

2. MASTER CAUTION light.

3. Grinding noise.

Procedure.

1. LAND AS SOON AS POSSIBLE.

If accompanied by C BOX OIL HOT caution
light and/or C BOX OIL PRESS caution light:

1. LAND IMMEDIATELY.

42 DEGREE AND 90 DEGREE GEARBOX
MALFUNCTIONS.

42 Degree/90 Degree Gearbox Oil
Overtemperature, or Low Pressure.

If GEARBOX OIL PRESS indicator is below
limits, or if combining GEARBOX OIL TEMP
is above limits:

1. LAND AS SOON AS POSSIBLE.

Indications.

1. 42°/90° TEMP/PRESS caution light.

2. MASTER CAUTION light.

Procedure.

1. Collective — DECREASE.

2. MASTER CAUTION light — RESET.

3. LAND AS SOON AS POSSIBLE.

42°/90° Chip Detector
Indications.

1. 42°/90° CHIP DETR caution light.

2. MASTER CAUTION light.

Procedure.

1. Collective — DECREASE.

2. MASTER CAUTION light - RESET.

3. LAND AS SOON AS PRACTICAL.

If accompanied by 42°/90° TEMP/PRESS caution
light:

1. LAND AS SOON AS POSSIBLE.

ROTOR BRAKE PRESSURIZED IN
FLIGHT.

Indications.

1. ROTOR BRAKE warning light.

2. Rotor brake handle out of down position.

3. Decrease in rotor rpm.

4. RPM caution light and audio.

Procedure.

1. Rotor brake handle — FULL

DOWN.

If warning light remains on:

2. HYP switch — SYS 2 OFF.

3. ROTOR RPM — MONITOR.

CAUTION

<; <;

With hydraulic system 2 OFF, SC AS
pitch and roll channels will be
inoperative.

4. LAND AS SOON AS PRACTICAL.

If warning light remains on:

1. LAND AS SOON AS POSSIBLE.

WING STORES JETTISON.

Each of the four ejector racks are equipped with an
electrically operated ballistic device to jettison the
attached weapon. The pilot can select one or all
stations for jettison. The copilot/gunner can select
inboard, outboard or both.

NOTE

Jettisoning inboard stores with 4 TOW
launchers installed will cause the out¬
board store to be jettisoned first regard¬
less of outboard jettison switch positions.

5-28

Change 1

Pilot Procedures For Jettisoning.

1. EMERGENCY JETTISON SELECT - ON
(as appropriate).

2. JETTISON button - DEPRESS (at least one
second).

3. EMERGENCY JETTISON SELECT — AS
APPROPRIATE.

Copilot/Gunner Procedures for Jettisoning.

1. WING STORES JETTISON INBD, OUTBD,
or BOTH (as appropriate).

2. JETTISON cover — UP.

3. WGST JTSN — UP.

4. JETTISON cover — DOWN.

5. WING STORES JETTISON — BOTH.

LOST PLANE PROCEDURES.

The primary requirements when lost are as
follows:

1. Confess.

2. Climb.

3. Conserve.

4. Communicate.

5. Conform.

LOST SIGHT DURING IMC

In the event of lost sight during IMC flight, the
reversal base course will be the reciprocal of the
flight present heading. Upon signal, the helicopters
will acknowledge and take the following action.
See figure 5-2.

1. Helicopter numbers two and four will
commence a standard rate turn away from
the flight. They will call passing through 90
degrees of turn and will turn 170 degrees.

2. Helicopter number three will climb 500 feet on
the present heading, after completing the
climb the helicopter will reverse heading
away from the flight leader, 170 degrees.
When number four helicopter reports passing

through 90 degrees upon completing the
reversal turn, helicopter number three will
descend to the initial altitude.

3. The flight leader, upon receiving the radio
call of helicopter number two passing
through 90 degrees of turn, will reverse course

60 SEC - 4 -

90 SEC

60 SEC

90 SEC

180°

30 SEC

- RADIO

/ CALL

-170°

STANDARD RATE TURNS.

CLIMB AND DESCEND AT 500 fpm.
MAINTAIN DISPERSAL AIRSPEED.
RENDEZVOUS WHEN VFR IS REGAINED.

180 degrees on the same side as helicopter
number two.

4. It is essential that all helicopters maintain
the airspeed of the flight when the dispersal
was commenced. The flight will regroup
when in a clear area.

90 SEC

60 SEC

-4-120 SEC

\ DESCEND
\ 500’

30 SEC

RADIO \

CALL \

170°

Figure 5-2. Lost Sight During IMG Flight Procedures

60 SEC

-V 90 SEC

210900-133

5-29

Section V
Part 4

NAVAIR 01 -HIAAB-1

PART 4 — LANDING EMERGENCIES

When anticipating an emergency
landing or ditching, each crewman
should place his shoulders against the
seat back, manually lock the shoulder
harness and keep back straight to obtain
maximum protection from the restraint
system.

AUTOROTATIVE LANDING.

ground, the cyclic stick should be moved slightly
forward of the neutral position. After touchdown,
decrease collective slowly to full down.

NOTE

The best glide airspeed is 110 KIAS. The
minimum rate of descent airspeed is 65
KIAS.

Procedures.

NOTE

A safe autorotative approach and landing is
dependent upon variables such as pilot capability,
density altitude, airspeed, gross weight, proximity
of suitable landing area, plus wind direction and
velocity. This does not preclude operation in the
restricted height velocity area during emergencies
or pressing operational requirements. Heading is
maintained by applying right pedal to decrease
the tail rotor thrust. Autorotative rotor rpm will
vary with ambient temperature, pressure altitude,
G loading, and gross weight. High gross
weights, increased G loads, and higher altitudes
and temperature will cause increased rotor rpm
which can be controlled by increasing collective.
Do not exceed 120 KIAS in sustained autorotation.

NOTE

Avoid abrupt control movement during
high speed autorotation to prevent over
controlling.

Any increase of rotor rpm, above that specified for
maximum glide, will result in increased rate-of-
descent. At an altitude of 100 to 75 feet, a flare
should be established by moving the cyclic stick
aft. This will decrease both airspeed and rate-of-
descent and cause an increase in rotor rpm that is
dependent upon the rate that the flare is executed.
Increased rotor rpm is desirable because more
energy will then be available to the main rotor
when collective is applied. Sites for autorotative
landings should be hard, flat, smooth surfaces
clear of approach and rollout obstructions. During
landing the helicopter should be held in a skids
level attitude and, when contact is made with the

If time and altitude permit, engine
airstart may be attempted after engine
failure. It is usually better to concentrate
on making a safe landing than to use
valuable time attempting an airstart.

1. Autorotation — ESTABLISH.

NOTE

All autorotative landings should be
made into the wind if possible.

2. Throttles — CLOSE.

3. FUEL — OFF.

4. Cyclic — AS REQUIRED (to reduce rate-of-
descent and airspeed).

5. Collective — INCREASE (as required to
complete landing).

After touchdown:

6. Collective — FULL DECREASE.

7. Rotor brake — ENGAGE.

8. Helicopter — SHUTDOWN.

SINGLE ENGINE LANDING

Procedure.

Under certain conditions, airspeed in excess of
25 KIAS may be necessary to land single engine.

5-30

NAVAIR 01 -HIAAB-1

Section V
Part 4

LANDING IN TREES.

When rotor blades have stopped:
6 . Helicopter — EXIT.

An autorotation into a heavily wooded area should
be accomplished by executing a normal

autorotative approach and full flare. The flare when well clear of the helicopter,
should be executed so as to reach zero rate of

descent and zero ground speed as close to the top of 7 L if e vest — INFLATE,

the trees as possible. As the helicopter settles,
increase collective to maximum.

DITCHING.

When the decision is made to ditch:

• I) o not abandon helicopter until rotor
blades have stopped.

• Do not inflate life vest until well clear of
the helicopter.

Procedure.

1 . Transponder MASTER — EMER.

2. MAYDAY — BROADCAST (give position).

Ditching — Power On.

Procedure.

Perform normal approach to hover 3 to 5 feet above

the water.

1. Canopy jettison handle — ROTATE
(counterclockwise 90 degrees) PULL.

2. Both throttles — CLOSE.

3. Collective — INCREASE (smoothly to
cushion landing).

As helicopter settles:

4. Collective — INCREASE (to maximum).

5. Rotor brake — ENGAGE.

Ditching — Power Off.
Procedure.

1 . Autorotation — ESTABLISH.

2. Canopy jettison handle — ROTATE
(counterclockwise 90 degrees) PULL.

Helmet visors shall be down prior to
activation of the CRS to preclude
eye injury.

3. Collective — INCREASE (smoothly to
cushion landing).

As helicopter settles:

4. Collective — INCREASE (to maximum).

5. Rotor brake — ENGAGE.

When rotor blades have stopped:

6 . Helicopter — EXIT.

When well clear of helicopter:

7. Life vest — INFLATE.

5-31/(5-32 blank)

NAVAIR 01-H1AAB-1

Section VI
Part 1

SECTION VI—ALL WEATHER OPERATION

TABLE OF CONTENTS

Introduction.6-1

PART 1 — INSTRUMENT PROCEDURES

Instrument Flight Procedures.6-1

PART 2 - EXTREME WEATHER OPERATION

Cold Weather Operation.6-2

Hot Weather Operation.6-6

Mountain and Rough Terrain Flying.6-6

INTRODUCTION.

The purpose of this section is to provide information
and procedures for operating under light icing, cold
weather and instrument flight conditions. This
section does not include equipment descriptions since
this information is contained in Section I. Detailed
instrument procedures are discussed in the
NATOPS INSTRUMENT FLIGHT MANUAL.

NOTE

• Because of various controllable modes of
helicopter flight, the possibility of pilot
vertigo caused by sideward motion or
oscillation is a more prevalent hazard

during night and instrument flight than it
is in fixed-wing flight.

• Under instrument conditions
particularly at night, through
conditions of reduced visibility,
unnecessary operation of the anti¬
collision light should be avoided.
Uncommon reflection on the helicopter’s
windows caused by rotating light being
reflected back from the clouds through
the whirling blades may cause vertigo.
Crew coordination is discussed in
Section IX.

PART 1 — INSTRUMENT PROCEDURES

INSTRUMENT FLIGHT PROCEDURES.

Simulated Instrument Flight.

Safety precautions and detailed procedures for
conducting simulated instrument flights are
contained in OPNAVINST 3710.24 series.

Start.

Complete normal exterior inspections and the
prestart/start checklist items.

Instrument Flight Checklist.

1. Maps, supplement, approach plates — As
required.

6. Magnetic compass — Check.

7. Vertical velocity indicator — Check needle
position.

8. Altimeters — Check and set.

9. Clock — Set.

10. Radios and IFF — Check and set.

11. AN/ASN-75 compass — Slaved, check
alignment.

Air Taxi.

2. Fuel packet — If required.

3. Cockpit heating equipment — Check operation.

4. PITOT heater — Check operation.

5. RAIN RMV — Check operation.

CAUTION

Extended use of the RAIN RMV system
can cause damage.

1. AN/ASN-75 compass — Check operation.

2. Turn and slip indicator — Check alignment and
operation.

3. Attitude indicator — Check alignment,
operation, and set horizontal bar.

4. Magnetic compass — Check operation.

5. Vertical velocity indicator — Check operation.

6. Exterior lights — As desired.

6-1

Section VI
Part 1 - Part 2

NAVAIR 01 -HI AAB-1

Instrument Takeoff.

When a normal hover is not possible, the helicopter
may be flown off the deck and into a normal climb
without any outside reference.

1. Maintain a level attitude with reference to the
attitude indicator.

2. As the helicopter becomes airborne, move the
cyclic control stick forward and adjust collective
pitch as necessary for transition into a forward
speed climbing flight.

NOTE

The airspeed indicator is unreliable at
airspeeds less than 40 knots.

3. Establish a rate-of-climb of at least 500 feet
per-minute with reference to the altimeter and
vertical speed indicator.

NOTE

Normally, turns should not be executed
prior to reaching 200 feet altitude.

4. Maintain a smooth acceleration up to 100
knots with reference to the attitude indicator
and the airspeed indicator.

Instrument Climb.

Climb under instrument conditions is similar to the
climb technique and procedure described in Section
HI. Under instrument conditions use the best rate of
climb speed for the operating gross weight. Climbing

turns should be limited to a maximum bank of 20
degrees.

Instrument Cruising Flight.

After leveling off, stabilize airspeed and power.
Particular attention should be given to navigation
since the slow airspeed associated with helicopters
can result in large drift angles.

SPEED RANGE.

A minimum speed of 70 knots should be observed
to maintain the normal flight characteristics
associated with forward flight.

ELECTRONIC EQUIPMENT.

Radio and navigation equipment are operated in the
normal manner.

HOLDING.

An airspeed of approximately 100 knots can be
easily maintained in the normal holding pattern.
However, a navigational problem will be present
while attempting to maintain a pattern in high
wind.

NOTE

Drift correction angles of 30 degrees are
not uncommon to a helicopter.

Descent.

Normal descents are made by reducing power until
the desired rate-of-descent is accomplished. Enroute
descents are normally made at cruising airspeed.

PART 2 - EXTREME WEATHER OPERATION

COLD WEATHER OPERATION.

Introduction.

Operation of the helicopter in cold weather or an
arctic environment presents no unusual problems
if the pilot is aw'are of the changes that take place
and conditions that may exist because of the lower
temperatures and freezing moisture. The pilot
must be more thorough in the walk-around
inspection when temperatures have been or are
below zero degrees C (plus 32 degrees F).

Engine Servicing.

Fuel and oil servicing should be accomplished
immediately after engine shutdown to prevent
condensation within the tanks due to temperature
change. Refer to the Servicing Diagram figure 1-
32.

Engine Ground Operation.

During extreme conditions, install covers after
engine shutdown. In extreme cold weather ground
heater unit may be used.

6-2

NAVAIR 01-H1AAB-1

Section VI
Part 2

Snow, slush, or ice shall be removed from any area
where jet engines may be operated. Keeping the
areas clean will prevent cinders, sand, or chunks of
ice from being sucked into the engines or blown at
high velocity into other aircraft that might be in the
vicinity.

During extreme cold weather, external vents and
drains shall be inspected prior to operating engines
and prior to flight.

I |

CAUTION

Should the engines fail to accelerate to
proper idle speed (cold hang-up) or the
time from light-off to idle is excessive,
abort start. Refer to utilization of Manual
Fuel for Cold Start Section VI.

An auxiliary power unit should be used, when
available, to ensure a smooth, fast engine
acceleration.

A sudden loss of oil pressure in cold weather, other
than a drop caused by relief valve opening, is
usually due to a broken oil line. Shutdown and
investigate for cause.

Install engine inlet and exhaust covers after
shutdown.

Utililization Of Manual Fuel For Cold Start.

Certain characteristics of the fuel control during
cold weather operation, although not affecting
the engine light off capabilities, will cause the gas
producer turbine to hangup and not accelerate to
idle speed, with Ng and ITT remaining low. In this
case, the GOV switch may be placed in MANUAL
after light off to bypass the automatic features
of the control and provide manual scheduling of
fuel as selected by the pilot.

Rapid throttle movement with possible ice
accumulation in the compressor could result in
stall when GOV switch is positioned to MANUAL.
Refer to Underspeeding Nf Governor, Section V.
After start has been completed, the GOV switch
should be switched to AUTO.

CAUTION

Do not operate the engine in excess of 71
percent Ng until engine and combining
gearbox oil temperatures reach +15
degrees C.

Preparation for Flight.

Preparation for cold weather flights should include
normal procedures in Section IV with the following
exceptions or additions: All vents and openings such
as fuel vents, battery vents, transmission breather,
heater exhaust and intake, and engine air intakes
must be checked for ice.

Accumulation of snow and ice will be
removed prior to flight. Failure to do so
can result in hazardous flight, due to
aerodynamic and center of gravity
disturbances as well as the introduction of
snow, water and ice into internal moving
parts and electrical systems. The pilot
should be particularly attentive to the
main and tail rotor systems and their
exposed control linkages.

NOTE

At temperatures of minus 35 degrees C
(minus 31 degrees F) and lower, the grease
in the couplings of the main transmission
driveshaft may congeal to a point that the
couplings cannot operate properly. If
found frozen, apply heat to thaw the
couplings, before attempting to start the
engines. Indication of proper operation is
obtained by turning the main rotor blade
opposite to the direction of the rotation
while an observer watches the driveshaft
to see that there is no tendency for the
transmission to "wobble” while the
driveshaft is turning.

PREHEATING.

Whenever outside ambient temperatures are ■
minus 40 degrees, preheating of the engines, I
gearbox, transmission, and associated system M
components is required. Flight and engine controls |
may be difficult to move after the helicopter has
been cold soaked. If the controls are not
sufficiently free for a safe start and low power
warm-up, have the affected controls thawed by
heating. It may also be advisable to apply pre¬
heating to other areas such as the engines, trans¬
mission, main rotor hub, and cockpit.

NOTE

When moving the helicopter into or out of
a heated hangar where there is an extreme

Change 1

6-3

Section VI
Part 2

NAVAIR 01-H1AAB-1

difference in outside temperature, a
canopy door should be open slightly to
equalize the temperature inside the
cockpit. Extremely unequal temperatures
on opposite sides of plexiglas can cause
differential contraction and breakage.

Main Rotor Blades and Elevator.

Visually check upper surfaces to be free of ice and
snow. Untie the blades and walk through 360 degrees
in the direction of rotation and check to see that there
is no restriction in operation or flapping freedom due
to ice formation. Check synchronized elevator for ice
and snow on surface and for restricted movement due
to ice and snow between fuselage and elevator.

Before Starting Engines.

An auxiliary power unit should be used when
available to ensure a smooth, fast engine
acceleration to preclude a hot start.

CAUTION

Whenever possible, avoid starting engines
on glare ice to avoid the effect of torque
reaction when increasing rpm.

NOTE

Battery starts below minus 15 degrees
C are marginal.

Starting Engines.

When outside air temperature is between minus 18
degrees C and minus 32 degrees C (zero degrees F and
minus 25 degrees F), accomplish the following
procedures in addition to those listed in Section II.

Do not advance beyond 71 percent Ng until both
engines, combining gearbox, and transmission oil
pressures are stabilized within desired operating
range.

Under cold weather conditions, make sure
all instruments have warmed up
sufficiently to ensure normal operation.
Check for sluggish instruments before
takeoff.

CAUTION

A sudden loss of oil pressure in cold
weather, other than a drop caused by relief
valve opening, is usually due to a broken
oil line. Shutdown and investigate for
cause.

NOTE

Before takeoff under icy conditions, check
that landing gear is not frozen to the
ground.

ICING.

Icing in the air intake system will be evidenced in the
cockpit only as a power loss (which could be as much
as 5 percent) and a corresponding increase in Ng,
RPM of less than 2 percent.

Air Taxi (Snow Conditions).

Operating the helicopter during conditions of snow
may result in a hazardous situation known as
“whiteout”. Whiteout, or circulation of snow
through the rotorwash, can occur during air taxi,
when in a hover, or on short final to landing. It is
potentially hazardous because the pilot may lose
visual reference outside the cockpit. Air taxiing
should be at an airspeed that will keep the snow
cloud aft of the stub wings (approximately 10-15
knots depending on the wing). Care should be
taken not to air taxi near another operating air¬
craft.

Takeoff.

NOTE

It may be necessary to get the aircraft
light on the skids and apply small pedal
pressure to ensure the skids are not
frozen to the ground.

Cold weather presents no particular takeoff problem
unless the cold weather is accompanied by snow.
The problem of restricted visibility due to blowing
or swirling snow (from rotor wash) can be acute
and may require a maximum power takeoff, or
perhaps even an instrument takeoff without hover
to get the helicopter safely airborne. Use available
objects for reference, such as smoke grenades, oil
drums, rocks, etc. If the takeoff is surrounded by
a large expanse of smooth, unbroken snow there
is danger that the pilot may become disoriented
because of the absence of visible ground reference
objects.

Icing Conditions.

Flight through icing conditions should be avoided.
However, should icing be inadvertently encountered,
the aircraft may exhibit the following
characteristics:

ENGINES.

The engines will continue to perform with no
noticeable power degradation for up to 60 minutes in
light icing (V 2 inch per 40 miles) at temperatures
down to -20 degrees C.

6-4 Change 1

NAVAIR 01-H1AAB-1

Section Vi
Part 2

RAIN REMOVAL AND PITOT HEAT.

Both of these systems will continue to perform
satisfactorily in icing conditions. Both systems
should be activated when flying through visible
moisture at temperatures of 0 degrees C and below.

ROTOR SYSTEM.

Ice will build on the rotor blades and shed naturally
in 15-20 minute cycles after encountering icing
conditions. Ice build-ups on the rotor system are
characterized by a concurrent increase in torque and
ITT. Initiation of the shed cycle usually results in a
one-per-rev vibration due to asymmetrical shedding.
Vibrations are of light to heavy intensity. If
vibrations are encountered immediately after or
during flight through icing conditions, rapidly
beeping the rotor system through its entire range
may alleviate the problem. Regardless, the shed cycle
should be complete in 5 to 10 minutes after vibrations
are initially encountered. After the shed cycle is
complete, torque and ITT should drop.

AIRFRAME.

On all snow landings anticipate the worst conditions;
that is, restricted visibility due to loose whirling
snow and an unfirm ice crust under the snow. When
loose or powdery snow is expected, make an approach
and landing with little or no hover to minimize the
effect of rotor wash on the snow. If possible, have
some prominent ground reference objects in view
during the approach and landing. If no such objects
are available, a smoke grenade, etc. dropped from the
helicopter may suffice.

In flights of two or more, separation should be
extended prior to arriving in the landing zone to
preclude the possibility of having to land in a
snow cloud produced by another aircraft.

If visual reference is lost, accomplish a
go-around.

CAUTION

Ice build-up can be monitored by observing the stub
wing and the forward section of either skid.

IN-FLIGHT.

Prior to or immediately after encountering icing
conditions, use RAIN RMV to keep windshield
clear prior to ice formation.

<; ;,

CAUTION

<; ;!

• Whenever possible, when landing on glare
ice, reduce sink rate as much as practical,
in order to reduce bending loads on the
crosstubes.

$ Radio and radar waves can penetrate the
surface of snow and ice fields, (such as the
polar region) therefore when radio and
radar equipment are used for measuring
terrain clearance, they may indicate
greater terrain clearance than actually
exists.

RAIN RMV system should be turned
OFF as soon as cleared vision will
permit. Heat may melt windshield if
operated for a lengthy period on a dry
windshield.

Landing.

In normal operations helicopters are often required
to land or maneuver in areas other than prepared
airfields. In cold weather this frequently involves
landing and taking off from snow covered terrain.
The snow depth is usually less in open areas where
there is little or no drift effect. The snow depth is
usually greater on the downwind side of ridges and
wooded areas. Whenever possible, the pilot should
familiarize himself with the type of terrain under the
snow (tundra, brush, marshland, etc.).

After contacting the surface, maintain rotor rpm and
slowly decrease collective pitch, while slightly
rotating the cyclic stick until the helicopter is firmly
on the ground. Be ready to takeoff immediately if,
while decreasing collective pitch, one landing gear
should hang up or break through the crust. Do not
reduce rotor rpm until it is positively determined
that the helicopter will not settle.

Shutdown.

The rotor brake should not be used on shutdown to
preclude airframe damage due to inducing main
rotor blade ice shedding. Should operational
requirements dictate use of the rotor brake, the rotor
system should be allowed to coast down to 30 percent
RPM before gently applying the rotor brake.

Change 1

6-5

NAVAIR 01 -HIAAB-1

Section VI
Part 2

Post-flight.

Inspect underside of both main rotor blades after
shutdown for possible damage due to tail rotor blade
ice shedding.

Before Leaving The Helicopter.

Perform the following checks in addition to those
listed in Section II: Open pilot and copilot/gunner
canopy doors approximately one and one-half inches
to permit free circulation of air to retard frost
formation and reduce cracking of transparent
surfaces due to differential contraction. Check that
moisture accumulations are drained as soon as
possible after engine shutdown. Check fuel cell
sumps, fuel strainer, transmission oil sump, and
engine oil systems. Check all vents for ice stoppage.

HOT WEATHER OPERATION.

Operations when outside air temperatures are above
standard day conditions do not require any special
handling technique or procedures, other than a closer
monitoring of oil temperatures and ITT. As ambient
temperature increases, engine efficiency decreases;
and power can become critical under high gross
weight conditions on extremely hot days.

Desert Operation.

Desert operation generally means operation in a very
hot, dusty, and often windy atmosphere. Under such
conditions sand and dust will often be found in vital
areas of the helicopter. Severe damage to the affected
parts may be caused by sand and dust. The helicopter
should be towed into takeoff position, which if at all
possible, should be on a hard, clear surface, free from
sand and dust. Ensure the engine inlets are free of
sand, heavy dust accumulation, and other foreign
matter. Use normal starting procedures.

NOTE

During warm weather, oil temperature
will probably be on the high side of the
operating range.

Install engine inlet and exhaust covers after
shutdown.

Preparation For Flight.

Plan the flight thoroughly to compensate for existing
conditions by using the charts in Section XI. Check

for the presence of sand and dust in control hinges
and actuating linkages. Inspect for, and have
removed, any sand or dust deposits on instrument
panel and switches, and on and around flight and
engine controls.

MOUNTAIN AND ROUGH TERRAIN
FLYING.

Many helicopter missions require flight and landings
in rough and mountainous terrain. Refined flying
techniques, along with complete and precise
knowledge of the individual problems to be
encountered, are required. Landing site condition,
wind direction and velocity, gross weight
limitations, and effects of obstacles are a few of the
considerations for each landing or takeoff. In a great
many cases, meteorology facilities and information
are not available at the site of intended operation.
The effects of mountains and vegetation can greatly
vary wind conditions and temperatures. For this
reason, each landing site must be evaluated at the
time of intended operation. Altitude and
temperature are major factors in determining
helicopter power performance. Gross weight
limitations under specific conditions can be
computed from the performance data in Section XI. A
major factor improving helicopter lifting
performance is wind. Weight carrying capability
increases rapidly with increases in wind velocity
relative to rotor system. However, accurate wind
information is more difficult to obtain and more
variable than other planning data. It is therefore not
advisable to include wind in advanced planning data,
except to note that any wind encountered in the
operating area may serve to improve helicopter
performance. In a few cases, operational necessity
will require landing on a prepared surface at an
altitude above the hovering capability of the
helicopter. In these cases, a sliding landing and
takeoff will be necessary to accomplish the mission.
Data for these conditions can be computed from the
charts in Section XI.

Wind Direction and Velocity.

There are several methods of determining the wind
direction and velocity in rough areas. The most
reliable method is by the use of smoke generators.
However, it must be noted that the hand held
day/night distress signal and the standard ordnance
issue smoke hand grenade, although satisfactory for
wind indication, constitute a fire hazard when used
in areas covered with combustible vegetation.
Observation of foliage will indicate to some degree

6-6

NAVAIR 01 -HIAAB-1

Section VI
Part 2

the direction of the wind, but is of limited value in
estimating wind velocity. Helicopter drift
determined by eyesight without the use of
navigational aids is the first method generally used
by experienced pilots. The accuracy with which wind
direction may be determined through the "drift”
method becomes a function of wind velocity. The
greater the wind value the more closely the direction
may be defined.

Landing Site Evaluation.

Five major considerations in evaluating the landing
area are: (1) height of obstacles which determine
approach angle; (2) size and topography of the
landing zone; (3) possible loss of wind effect; (4) power
available; and (5) departure route. The transition
period is the most difficult part of any approach. The
transition period becomes more critical with
increased density altitude and/or gross weight,
therefore approaches must be shallower and
transition more gradual. As the height of the
obstacles increase, larger landing areas will be
required. As wind velocity increases so does
helicopter performance; however, when the
helicopter drops below an obstacle a loss of wind
generally occurs as a result of the airflow being
unable to immediately negotiate the change
prevalent at the upwind side of the landing zone
where a virtual null area exists. This null area
extends toward the downwind side of the clearing and
will become larger as the height of the obstacle and
wind velocity increases. It is therefore increasingly
important in the landing phase that this null area be
avoided if marginal performance capabilities are
anticipated. The null area is of particular concern in
making a takeoff from a confined area. Under heavy
load or limited power conditions it is desirable to
have sufficient airspeed and translational lift prior to
transitioning to a climb, so that the overall climb
performance of the helicopter will be improved. If the
takeoff cycle is not commenced from the most down
wind portion of the area and translational velocity
achieved prior to arrival in the null area, a
significant loss in lift may occur at the most critical
portion of the takeoff. It must also be noted that in the
vicinity of the null area nearly vertical downdraft of
air may be encountered, which will further reduce
the actual climb rate of the helicopter. It is feasible
that under certain combinations of limited area, high
obstacles upwind, and limited power available, the
best takeoff route would be either crosswind or
downwind, terrain permitting. The effects of
detrimental wind flow and the requirement to climb

may thus be minimized or circumvented. Even
though this is a departure from the cardinal rule of
"takeoff into the wind” it may well be the proper
solution when all factors are weighed in their true
perspective. Never plan an approach to a confined
area wherin there is no reasonable route of
departure. The terrain within a site is considered
from an evaluation of vegetation, surface
characteristics, and slope. Care must be taken to
avoid placing the rotor in low brush or branches.
Obstacles covered by grass may be located by
flattening the grass with rotor wash prior to landing.
Power should be maintained so that an immediate
takeoff may be accomplished should the helicopter
start tipping from soft earth or a skid being placed in
a hidden hole.

Effects of High Altitude.

Engine power available at altitude is less, and
hovering ability can be limited. High gross weight at
altitude increases the susceptibility of the helicopter
to blade stall. Conditions that contribute to blade
stall are high forward speed, high gross weight, high
altitude, induced "G” loading and turbulence.
Shallower turns at slower airspeeds are required to
avoid blade stall. A permissible maneuver at sea
level must be tempered at a higher altitude. Smooth
and timely control application and anticipation of
power requirements will do more than anything else
to improve altitude performance.

Turbulent Air Flight Techniques.

Helicopter pilots must be constantly alert to evaluate
and avoid areas of severe turbulence; however, if
encountered, immediate steps must be taken to avoid
continued flight through it to preclude the structural
limits of the helicopter being exceeded. Severe
turbulence is often found in thunderstorms, and
helicopter operations should not be conducted in
their vicinity. The most frequently encountered type
of turbulence is orographic turbulence. It can be
dangerous if severe and is normally associated with
updrafts and downdrafts. It is created by moving air
being lifted by natural or man made obstructions. It
is most prevalent in mountainous regions and is
always present in mountains if there is a surface
wind. Orographic turbulence is directly proportional
to the wind velocity. It is found on the upwind of
slopes and ridges near the tops and extending down
the downwind slope (figure 6-1). It will always be
found on the tops of ridges associated with updraft in
the upwind side and downdrafts on the downwind

6-7

Section VI
Part 2

NAVAIR 01 -HI AAB-1

side. Its extent on the downwind slope depends on the
strength of the wind and the steepness of the slope. If
the wind is fairly strong (15 to 20 knots) and the slope
is steep, the wind will have a tendency to blow off the
slope and not follow it down; however, there will still
be some tendency to follow the slope. In this situation
there will probably be severe turbulence several
hundred yards downwind of the ridge at a level just
below the top. Under certain atmospheric conditions,
a cloud may be observed at this point. On more gentle
slopes the turbulence will follow down the slope, but
will be more severe near the top. Orographic
turbulence will be affected by other factors. The
intensity will not be as great when climbing a smooth
surface as when climbing a rough surface. It will not
follow sharp contours as readily as gentle contours.
Man-made obstructions and vegetation will also
cause turbulence. Extreme care should be taken
when hovering near buildings, hangers, and similar
obstructions. The best method of overfly ridge lines
from any direction is to acquire sufficient altitude
prior to crossing to avoid leeside downdrafts. If
landing on ridge lines (figure 6-2), the approach
should be made along the ridge in the updraft, or
select an approach angle into the wind that is above
the leeside turbulence. When the wind blows across a
narrow canyon or gorge (figure 6-3) it will often veer
down into the canyon. Turbulence will be found near
the middle and downwind side of the canyon or gorge.
When a helicopter is being operated at or near its
service ceiling and a downdraft of more than 1.6 feet
per second is encountered the helicopter will descend.

Although the downdraft does not continue to the
ground, a rate-of-descent may be established of such
magnitude that the helicopter will continue
descending and crash even though the helicopter is
no longer affected by the downdraft. Therefore, the
procedure for transiting a mountain pass shall be to
fly close abeam that side of the pass or canyon which
affords an upslope wind. This procedure not only
provides additional lift but also provides a readily
available means of exit in case of emergency.
Maximum turning space is available and a turn into
the wind is also a turn to lower terrain. The often
used procedure of flying through the middle of a pass
to avoid mountains invites disaster. This is
frequently the area of greatest turbulence (figure
6-4) and in case of emergency, the pilot has little or no
opportunity to turn back due to insufficient turning
space. Rising air currents created by surface heating
causes convective turbulence. This is most prevalent
over bare areas. Convective turbulence is normally
found at a relative low height above the terrain,
generally below 2000 feet. It may, however, reach as
high as 8000 feet above the terrain. Attempting to fly
over convective turbulence should be carefully
considered, depending on the mission assigned. The
best method is to fly at ihe lowest altitude consistent
with safety. Attempt to keep your flight path over
areas covered with vegetation. Turbulence can be
anticipated when transitioning from bare areas to
areas covered by vegetation or snow. Convective
turbulence seldom gets severe enough to cause
structural damage.

NULL AREA USUALLY FOUND
ON VERY CREST OF SLOPE

WIND

UPDRAFTS WILL EXTEND ABOVE THE SURFACE FARTHER THAN
THE TURBULENCE DEPENDING ON WIND SPEED

IN VERY STRONG WIND CONDITIONS AND/OR VERY STEEP
SLOPES THE TURBULENCE WILL BE FOUND FORWARD OF THE
SLOPE IN CLEAR AIR.

-04947-18

Figure 6-1. Wind Flow Over and Around Peaks

6-8

NAVAIR 01 -HI AAB-1

Section VI
Part 2

Adverse Weather Conditions.

When flying and around mountainous terrain under
adverse weather conditions, it should be remembered
that the possibility of inadvertent entry into clouds is
ever present.

Air currents are unpredictable and may cause cloud
formations to shift rapidly. Since depth perception is
poor with relation to distance from cloud formation
and to cloud movement, low hanging clouds and scud
should be given a wide berth at all times. In addition
to being well briefed, the pilot should carefully study
the route to be flown. A careful check of the helicopter
compass should be maintained in order to fly a true
heading if the occasion demands.

Summary.

The following guide lines are considered to be most
important for mountain and rough terrain flying:

1. Make a continuous check of wind direction and
estimated velocity.

2. Plan your approach so that an abort can be made
downhill and/or into the wind without climbing.

3. If the wind is relatively calm try to select a hill
or knoll for landing so as to take full advantage
of any possible wind affect.

4. When evaluating a landing site in non-combat
operations, execute as many fly-bys as
necessary with at least one high and one low
pass before conducting operations into a
strange landing area.

5. Evaluate the obstacles in the landing site and
consider possible null areas and routes of
departure (figure 6-5).

6. Landing site selection should not be based solely
on convenience but consideration should be
given to all relevant factors.

7. Determine ability to hover out of ground effect
prior to attempting a landing.

8. Watch for rpm surges during turbulent
conditions. Strong updrafts will cause rpm to
increase, whereas downdraft will cause rpm to
decrease.

9. Avoid flight in or near thunderstorms.

10. Give all cloud formations a wide berth.

11. Fly as smoothly as possible and avoid steep
turns.

12. Cross mountain peaks and ridges high enough
to stay out of downdrafts on the leeside of the
crest.

13. Avoid downdrafts'prevalent on leeward slopes.

14. Plan your flight to take advantage of the
updrafts on the windward slopes.

15. Whenever possible, approaches to ridges
should be along the ridge rather than
perpendicular.

16. Avoid high rates of descent when approaching
landing sites.

17. Know your route and brief well for flying in
these areas.

6-9

Section VI
Part 2

NAVAIR 01-H1AAB-1

I. APPROACH THE UPWIND SIDE PARALLEL TO. OR. AT AS
SLIGHT AN ANGLE AS POSSIBLE TO THE RIDGELINE. RATHER
THAN PERPENDICULAR TO THE RIDGELINE.

2. IF TERRAIN DOES NOT PERMIT A PARALLEL APPROACH MAKE
APPROACH AS STEEP AS SAFELY POSSIBLE TO AVOID LEE¬
WARD BURBLE AND DOWNDRAFT.

PLAN AN ABORT ROUTE.

WINDWARD

(UPDRAFT]

cmc

Figure 6-2. Crosswind Effect On Pinnacle Approach

Figure 6-3. Wind Effect Over Gorge Or Canyon

6-10

NAVAIR 01 -HI AAB-1

Section VI
Part 2

IN THIS AREA DUE TO VENTURI EFFECT. L >
7 . FXCESSIVE TURBULENCE NEAR BOTTOM. ^

:()4‘)47-5:

Figure 6-4.

Wind Effect In Valleys Or Canyons

Figure 6-5.

Wind Effect in Confined Area

6-11/(6-12 blank)

NAVAIR 01-H1AAB-1

Section VII

SECTION VII — COMMUNICATIONS —
NAVIGATION EQUIPMENT
AND PROCEDURES

TABLE OF CONTENTS

Communications — Navigation — Identification
System..7-1

COMMUNICATIONS — NAVIGATION —
IDENTIFICATION SYSTEM.

Introduction.

The communications-navigation equipment for
the AH-lT (TOW) helicopter include a
FM(AN/ARC-114A) radio coupled with a
TSEC/KY-28 voice security unit. The helicopter
has a UHF(AN/ARC-159(V)1) coupled with a
TSEC/KY-28 voice security unit, and the UHF
direction finder (AN/ARA-50). Helicopter
intercommunication is controlled with (AN/AIC-
18). The helicopter is equipped with identification
radar (AN/APX-72) radar beacon (AN/APN-
154(V), radar altimeter (AN/APN-171(V),
compass set (AN/ASN-75B), automatic direction
finder (AI)F) (AN/ARN-83) and TACAN
(AN/ARN-84(V)). The primary navigation
instrument is the bearing distance heading
indicator which responds to the ADF, UHF and
TACAN units.

For antennas and their location see figure 7-1. Only
the antennas which the pilot can visually check for
damage are shown.

AN/ARC-114A FM Radio Set.

The AN/ARC-114A provides two-way frequency
modulated (FM) narrow band voice communications,
with homing capability, in the frequency range of
30.00 to 75.95 MHz (figure 7-2). However, homing is
primarily in the 30.00 to 60.00 MHz range. The set
operates on 920 channels for a distance of
approximately 50 miles as limited by line of sight.
FM homing information is presented to the course
displacement pointer (bug) and signal strength on
the vertical displacement pointer on the pilot
Attitude Direction Indicator (ADI).

The FM facility in conjunction with FM homing
antenna, develops course deviation (steering)
signal, FM signal adequacy, and station passage
signals. Receiver RF signals are routed from
homing antenna to FM radio set which compares
two signals and develops a course deviation
signal. This course deviation signal is applied to
the ADI left-right steering pointer. In addition, RF
input to FM radio set is monitored, output or
monitoring circuit is applied to signal adequacy
flag, and station passage pointer on ADI. In the
event of power failure or loss of homing signal, a
red flag will appear on the ADI.

The course'displacement pointer (located at the
bottom of the ADI) indicates left-right deviation from
the homing station. Flying the helicopter to the
pointer will return the pointer to null (on course
heading). The vertical displacement pointer (located
on the left side of the ADI) indicates signal strength
and station passage. The pointer moves up with
increasing signal strength. The FM homing
information is displayed automatically when the
ARC-114A mode selector switch is placed in the
homing position. The TSEC/KY-28 voice security
equipment is used with the FM Radio Set to provide
secure two-way communications.

TSEC/KY-28 VOICE SECURITY SYSTEM
(VHF-FM).

With the KY-28 equipment installed, received audio
is controlled by the KY-28 control (figure 7-2). In
voice security operation, encoded signals picked up
by VHF communication antenna are applied to the
VHF receivers. In the radio set receiver, signals are
selected, amplified and converted to coded audio
signals. Coded audio signals are then applied to

7-1

Section VII

NAVAIR 01 -HIAAB-1

AN/ARA-50

AN/APX-72

N2/83

210900-129

Figure 7-1. Antenna Location

7-2 Change 1

NAVAIR 01 -HIAAB-1

Section VII

FREQUENCY*
SELECTOR

T/R GUARD

M T/ R NOISE ■ TONE/X^ ^

\ F M

VOICE SE' ./ 1

' r

\J/

; s";

; ;

XCVR

FM . "i

7r\ [

NOMENCLATURE
FUNCTION SELECTOR
OFF
T/R

T/R GUARD

HOMING

RETRAN

FUNCTION

Power off.

Receiver — ON;
Transmitter — Standby

Guard receiver — ON;
Transmitter — Standby;
Receiver — ON

NOTE

Reception on the guard receiver is
unaffected by frequencies selected for
normal communications.

NOTE

The guard frequency can be selected on
the main receiver.

Activates the homing mode and displays on attitude
indicator. May also be used for normal voice
communications. The communications antenna is
automatically selected when the transmitter is keyed.

Not used.

210077-77-1

Figure 7-2. F M Radio (Sheet 1 of 2)

7-3

Section VII

NAVAIR 01 -HI AAB-1

NOMENCLATURE FUNCTION

FREQUENCY SELECTORS - INDICATOR

Left

Select first two digits of desired frequency.

Right

Select third and fourth digits of desired frequency.

RCVR TEST Switch

When pressed, audible tone indicates proper receiver
performance.

SQUELCH SELECTOR

NOISE

Eliminates background noise in headsets.

OFF

Deactivate squelch.

TONE/X

Squelches background noise in headsets.

AUDIO Control

Adjust receiver volume.

KY-28

OFF/ON/RLY switch

OFF — Remove power from KY28

ON — Applies power to KY28.

RLY — Not used.

C/ P switch

C — Permits ciphered communication on the radio set.

P — Permits unciphered communication on the radio set.

Zeroize switch
(under stripper cover)

Aft — Allows normal operation.

Forward — Placed in the forward position during
emergency situations to neutralize and make inoperative
the associated KY-28 cipher equipment. Do not place the
zeroize switch in the forward position unless a crash or
capture is imminent.

210077-77-2

Figure 7-2. F M Radio (Sheet 2 of 2)

KY-28 Coder (VHF) for translation to clear voice selection. When VHF monitor switch setting is made,

audio. The resulting decoded audio output is applied audio is amplified in the ICS control. Audio signal

to impedance matching network. Impedance level is further adjusted by the ICS control volume

matching network loads and distributes decoded (VOL). The decoded audio signal is then routed to

audio to pilot and copilot/gunner ICS control for pilot or copilot/gunner headset.

7-4

NAVAIR 01 -HI AAB-1

Section VII

With the KY-28 installed, transmitted audio signals
are used in conjunction with the VHF facility to
provide secure, two-way voice communications. The
encoder portion of the KY-28 coder (VHF) translates
the microphone audio to coded voice for application to
the VHF transmitter. Secure audio signal from the
VHF radio transmitter is sent to ground or air station
KY-28 coder (VHF) for translation to clear voice.

AN/ARC-114A FM Radio Set Operating
Procedures.

1. FM XCVR circuit breaker — CLOSED.

2. VOICE SECURITY FM circuit breakers —
CLOSED.

BATTERY switch - ON.

4. ICS transmit select — FM.

5. ICS monitor FM — ON.

6. ICS VOL control - AS REQUIRED.

7. Function selector — AS REQUIRED.

8. Frequency selector — AS REQUIRED.

9. SQUELCH - AS REQUIRED.

10. AUDIO - ADJUST.

11. RCVR TEST - PRESS.

KY-28

1. RLY switch — ON.

2. C/P switch — AS REQUIRED.

AN/ARC-159 (V) 1 UHF Radio Set

The AN/ARC-159 Radio Set is a solid state UHF
transceiver that provides two-way

amplitude-modulated communication capability
(figure 7-3). The radio set permits transmitting and
receiving on any of 7,000 frequencies, spaced 25-KHz
apart. In addition, the radio set is capable of
wide-band data (secure) communication, guard
frequency reception, and direction finder (DF)
reception. The guard receiver module is a
self-contained fix tuned receiver, set to some

frequency between 238.000 and 248.000 MHz
(usually 243.000 MHz) and can operate
simultaneously with the main receiver. During DF
reception, the radio set receives RF signals from the
DF antenna and routes the demodulated low
frequencies to the DF amplifiers and number 1
pointer of BDH indicators. The control panel, located
at the front of the radio set, provides for frequency
selection. Any 1 of 20 preset frequency channels or
any one of the 7,000 frequencies within the range of
the radio set may be selected. The copilot/gunner
has a guarded EMER UHF switch mounted on the
canopy frame. Activating the switch preempts any
other setting and places the system at 243.000
MHz.

UHF Radio Set Operating Procedure.

1. ICS PLT and GUNNER circuit breakers —
CLOSED.

2. UHF XCVR circuit breaker — CLOSED.

3. VOICE SECURITY UHF circuit breakers —
CLOSED.

4. BATTERY Switch — ON.

5. ICS transmit select — UHF.

6. ICS monitor UHF — ON.

7. ICS VOL control — AS REQUIRED.

8. UHF function selector — BOTH.

9. UHF frequency selectors or PRESET — AS
REQUIRED. '

10. UHF VOL control — AS REQUIRED.

11. UHF SQUELCH - ON.

12. Depress and release UHF TONE — 1020 Hz
TONE IS HEARD.

13. Mode selector — AS REQUIRED.

UHF Direction Finder (AN/ARA-50).

The direction finder group AN/ARA-50 is controlled
by the function selector switch on the UHF control
panel in ADF position (figure 4-3). System operates
in the frequency range of 225 to 339.95 MHz. Short or

7-5

Section VII

NAVAIR 01 -HI AAB-1

Figure 7-3. UHF Radio (Sheet 1 of 3)

7-6

NAVAIR 01 -HI AAB-1

Section VII

NOMENCLATURE FUNCTION

DIM control

Adjusts light intensity of LED
frequency display.

LAMP TEST
control

VOL control

Displays (888888) when the switch

is pressed.

Adjusts level of audio signal.

100 MHz
and 10 MHz
selector

Selects and indicates the manual
100 MHz and 10 MHz frequency
increments in operation.

1 MHz
selector

Selects and indicates the manual

1 MHz frequency increment in
operation.

0.1 MHz
selector

Selects and indicates the manual
0.1 MHz frequency increment in
operation.

50 KHz and

25 KHz
selector

Selects and indicates the manual
50 KHz and 25 KHz frequency
increments in operation.

PRESET

READ

Chart

PRESET

Permits selecting one of 20 preset
channels with the preset channel
control and displays the channel on
the readout indicator by the 4th
and/or 5th lamp. In this position the
manual selectors are ineffective.

Permits the operator to read the
frequency of the selected preset
channel. In this position the preset
frequency is displayed on the
readout indicator.

Provides a semipermanent
reference for all preset operating
frequencies set on the memory
drum.

Selects any one of 20 preset
frequency channels when the
mode selector is in the PRESET
position.

TONE control With TONE control (depressed),
switch the transmitter will transmit a 1 020

Hz tone signal.

SQ—OFF Enables or disables main receiver

switch squelch.

Mode selector Determines frequency selection and

switch indicates frequency and/or channel

selection.

GUARD Shifts transceiver to guard channel

frequency, 243.000 MHz and
displays 243.000 MHz on readout.
Both the preset and manual
frequencies are ineffective.

MANUAL Permits manual selection of any one
of 7,000 frequency channels by use
of the manual frequency selectors.
The frequency selected is displayed
on the readout indicator. In this
position the preset channel selector
control is ineffective.

Function Select mode of operation,

selector switch

OFF Removes power from radio set.

MAIN Equipment is in a transceiver mode

of operation. In normal condition
(key switch not actuated) the
equipment is in the receive
condition.

BOTH Equipment is energized in the same

way as described for the MAIN
position. In addition, the guard
receiver is turned on.

ADF AN/ARA-50 direction finding

equipment associated with
AN/ARC-159 Radio Set becomes
operative. Both the main and
guard receivers are enabled.

21007 7-7 8-2A

Figure 7-3. UHF Radio (Sheet 2 of 3)

7-7

Section VII

NAVAIR 01 -HI AAB-1

KY28

NOMENCLATURE
OFF/ON/RLY switch

C/P switch

Zeroize switch
(under stripper cover)

FUNCTION

OFF — Remove power from KY28.

ON — Applies power to KY28.

RLY - Not used.

C — Permits ciphered communication on the
radio set.

P — Permits unciphered communication on the
radio set.

Aft — Allows normal operation.

Forward — Placed in the forward position during
emergency situations to neutralize and make inoperative
the associated KY-28 cipher equipment. Do not place
the zeroize switch in the forward position unless a
crash or capture is imminent.

210077-78 -3

Figure 7-3. UHF Radio (Sheet 3 of 3)

long range of the DF set is selected by positioning
the UHF-DF RANGE switch (figure 7-4) on the
center section of the instrument panel, to LONG
or SHORT. When the function selector switch is
at the ADF position and a frequency of 225 to
339.95 MHz is selected, the UHF receiver is
coupled to the direction finder group and bearing
information to the station emitting the signal,
and is presented on the No. 1 pointer of the BDH
indicator.

UHF Direction Finder Operating Procedure.

1. UHF DF circuit breakers - CLOSED.

2. .Frequency selector knobs — POSITION (to
desired frequency.)

3. Function selector — ADF.

4. DF RANGE - POSITION (to LONG or
SHORT as required.)

7-8

NAVAIR 01-H1AAB-1

Section VII

210077-73

Figure 7-4. UHF Direction Finder

7-9

Section VII

NAVAIR 01 -HI AAB-1

NOTE

Approaching the station, there may be less
bearing pointer oscillation if the range
select switch is positioned to SHORT.

5. Bearing information — Read from No. 1 pointer
on BDH indicator.

6. To secure the equipment — Position the
function selector switch to OFF.

AN/APN 154(V) Radar Beacon.

The Radar Beacon receiver-transmitter is installed
in the tailboom. It is used to provide range
information, so as to extend the tracking range of a
ground-based radar which is a part of an
aircraft-navigation system. The required range
information is obtained in the following manner:
X-band signals transmitted by the ground base radar
interrogate the AN/APN-154( V). The
AN/APN-154(V), in turn replies to the interrogation
by transmitting a reply pulse to provide the range
information for the aircraft-navigation system. The
AN/APN-154(V) will reply to either coded or
noncoded signals, the mode of operation being
selected by aircraft personnel.

RADAR BEACON CONTROL PANEL.

The radar beacon control panel is marked RADAR
BEACON and is installed in the pilot instrument
panel (figure 7-5). The control contains two
switches designated as PWR and MODE. The
PWR switch is a toggle switch that has three
positions: OFF, STBY, and PWR. The MODE
switch is a six-position rotary switch that controls
mode of operation in the AN/APN-154(V).

Radar Beacon Operating Procedures.

An indicating device is not required for operation of
the AN/APN- 154(V). The only requirement for
operation is the positioning of the POWER and
MODE switches. When it is necessary to operate the
equipment:

1. PWR — PWR.

2. MODE - AS REQUIRED.

NOTE

Five minutes is required for the
AN/APN-154(V) to warm up. The PWR
switch may be placed at STBY for warm
up. However, it is impossible for the set
to transmit in STBY position.

AN/ASN-75B Compass Set.

The AN/ASN-75B Compass set provides an accurate
indication of helicopter heading. The AN/ASN-75B is
capable of operating as a magnetically slaved
directional gyro in the SLAVED mode, or as a free
directional gyro with latitude correction in the FREE
gyro mode. Mode selection facilities are provided by a
toggle switch on the panel of Compass Set Control
C-8021/ASN-75B. The AN/ASN-75B heading output
is presented on the rotating compass card of the
bearing-distance-heading indicator (BDHI) on the
pilot and copilot/gunner instrument panel. The
LATITUDE degree selector and LATITUDE N-S
switch are located on Amplifier-Power Supply
AM-4606/ASN-75 in the radio compartment
and must be placed on the desired position before
flight.

COMPASS SET CONTROL (C-8021/ASN-75B).

A control panel located on the pilot right console is
used to set the AN/ASN-75 compass system (figure
7-6). A knob marked PUSH TO SET provides for
synchronization of the gyro to the magnetic heading
of the helicopter. The synchronization indicator
(annunciator) is a zero-center meter marked 4- on the
left and • on the right. When the indicator reads -I-,
AN/ASN-75 may be synchronized by pressing
the PUSH TO SET knob located on the
C-8021/ASN-75 and rotating it in the 4 direction
until the indicator centers. Similarly, when the
indicator reads • the knob should be rotated in the •
direction. This ensures that the heading presented on
the compass card of the BDHI is correct. The PUSH
TO SET knob is also used to set the desired aircraft
heading while operating in the FREE gyro mode.
Exercise care to prevent setting the compass 180
degrees out of phase; in this situation the
synchronization indicator will center; however,
synchronization pointer movement will be opposite
to the direction of rotation of the control knob. The
synchronization indicator continues to provide a
visual check on the slaving operation. The
AN/ASN-75 compass system receives power from the
26 volt ac bus and the 115 volt ac bus. Circuit
protection is provided by a circuit breaker in the
circuit breakers panel labeled GYRO CMPS and
CMPS IND.

7-10

NAVAIR 01 -HI AAB-1

Section VII

NOMENCLATURE

PWR switch

MODE switch

ACLS test switch

FUNCTION

STBY position — power to AN/APN-1 54(V) is turned on but no output is
obtainable from receiver/transmitter due to disabling of decoder.

POWER position — encoder is enabled permitting normal operation of
AN/APN-1 54(V).

OFF position — primary input power to set is turned off.

SINGLE position — set will reply to single pulse (non-coded)
interrogations.

DOUBLE position — (5 positions) the set will reply only to properly coded
double-pulse interrogations that correspond to code determined by setting
of Mode switch.

ACLS test switch not used (for auxiliary KA band equipment). 210

Figure 7-5. Radar Beacon

7-11

Section VII

NAVAIR 01 -HI AAB-1

NOMENCLATURE
PUSH TO SET knob
SLAVED mode

FREE gyro mode

FUNCTION

Provides synchronization of gyro to magnetic heading of helicopter.

Will produce an output from transmitter synchro which agrees with
Remote Compass Transmitter Type ML-1 magnetic heading.

Provides correction for earth's rate of drift by Amplifier-Power Supply
AM-4606/ASN-75.

210077-80

Figure 7-6. Compass Set

7-12

NAVAIR 01 -HI AAB-1

Section VII

NOTE

When the SLAVED mode is selected, the
AN ASN-75 will automatically slave to
the correct heading at a rate of 2.5 ±1.25
degrees per minute. The correct heading
may be immediately selected with the
PUSH TO SET knob as explained above.

Once the correct heading is selected, the
AN ASN-75 will maintain the correct
magnetic heading on the BDHI compass
card.

When operating in areas of high latitude or during
ship board operations, the gyro may be unslaved to
prevent unreliable readings.

Slaved Gyro Operating Procedures.

1. Allow approximately two minutes for gyro to
reach operating speed.

2. Mode - SLAVED.

3. PUSH TO SET knob — Synchronize gyro and
magnetic heading by pushing in on knob and
rotating until synchronizing indicator is
centered.

Free Gyro Operating Procedures.

1. Check LATITUDE selectors and N-S switch
before flight.

2. Allow approximately two minutes for gyro
to reach operating speed.

| 3. Mode - FREE.

Intercommunications System (AN/AIC-18).

The intercommunications system provides
interphone communication between the pilot and
copilot/gunner within the helicopter (figure 7-7). It
also provides integrating facilities for the
communication and electronic equipment. An
interphone disconnect is provided on each wing tip.
When the microphone at the right crew station is
keyed it connects the right station to the pilot
microphone circuit. When the microphone at the left
crew station is keyed it connects the left side station
to the copilot/gunner microphone circuit.
Subsequently, if a transmitter is keyed at the pilot or
copilot/gunner station, and the microphone at the

right or left crew station is keyed that station would
also be put on the air through the respective ICS
amplifier.

INTERCOMMUNICATION CONTROL PANEL.

An INTER control panel is installed in the left
side of the pilot instrument panel and one is
installed forward of the copilot/gunner seat. The
control panel contains seven monitor switches
(six of which are utilized), a HOT MIC switch,
CALL switch, and VOLume control. Trans¬
mitter selector switch is located in the lower
center section of control panel.

Associated with the AN/AIC-18 is the press-to-
talk switch located on top of the pilot and
copilot/gunner cyclic stock grip. It is a 4 position
switch but only the forward and aft position are
utilized. Pressing forward keys the radio, aft keys
the ICS. The copilot/gunner has a footswitch for
keying the intercommunications system or radio,
as selected by the rotary selector switch.

Interphone Operating Procedures.

1. INTER PLT and ICS gunner circuit
breakers — CLOSED.

2. Monitor — PULL AND ROTATE.

3. Transmitter selector — AS DESIRED.

4. VOLume control — ADJUST.

5. Press-to-talk — RADIO/ICS.

AN/APX-72 Identification Transponder.

The AIMS Transponder system is composed of a
transponder (RT859/APX-72), a transponder
control panel (C-16280(P)/APX), an indicator/
encoder (AAU-21A), an on board transponder
self-test set (TS-1843/APX), transponder com¬
puter (KIT-1 A/TSEC), and an antenna. The
system provides IFF identification in response to
coded interrogations from ground, seaborne or
airborne stations. In addition, the signal returned
from the IFF transponder can be used by the
interrogating station to determine range and
azimuth information. Modes 1 and 2 are used for
military control purposes. Mode 3/A is the

7-13

Section VII

NAVAIR 01 -HI AAB-1

Q? Of QE, Of «

o?o,eo 6?'

COPILOT/GUNNER
FOOT SWITCH

ICS

PLT

GUNNER

•ic \

PILOT AND COPILOT/GUNNER
CYCLIC STICK GRIP

210077 - 63-11

Figure 7-7. Intercommunications Systems (Sheet 1 of 2)

7-14

NAVAIR 01 -HI AAB-1

Section VII

NOMENCLATURE

FUNCTION

Monitor switches Monitor the communication they represent. Any combination or

INT, UHF, FM, TACAN, a* 1 of switches may be in on position at one time for monitoring.

ADF and IFF

HOT MIC switch

Enable operator to conduct hand free intercommunications.

CALL switch

Provides emergency intercommunications on the call audio line,
regardless of the position of switches on the other station.

Transmitter Selector switch
(Marked INT, UHF, FM)

VOL control
Press-to-Talk

Copilot/Gunner
Foot switch

Enables operation and modulation of UHF and FM transmitters
also provides audio monitoring and side-tone from the associated
receiver.

INT position permits intercommunication between pilot and
copilot, gunner. (The Press-to-Talk switch on the cyclic stick grip
must be in the ICS or radio position to talk.

Vary audio level in headset.

Switch forward position keys the selected transmitter or the
interphone system if INT is selected on transmitter selector
switch.

Aft position keys the interphone circuit regardless of position of
transmitter selector switch.

Used to key selected transmitters. Key intercommunications
system if transmitter selector switch is in INT position.

210077-63-2 A

Figure 7-7. Intercommunications Systems (Sheet 2 of 2)

common military/civilian mode and is used by the
FAA for air traffic control purposes. On heli¬
copters modified by AFC-80, mode C is the
altitude reporting mode. With the M-C switch in
the ON position, it provides automatic reporting
of pressure altitude to ground stations. This
information is provided by the pilot’s altitude
indicator/encoder. Mode 4 is the secure (en¬
crypted) military identification mode, and is
operational only when the computer KIT-1A/
TSEC is installed.

The transponder, control panel, set-test and
altimeter vibrators receive 28 volt dc power from
the dc essential bus. The transponder and control
panel are protected by the IFF XPONDER circuit
breaker. The self-test set is protected by an
adjacent circuit breaker labeled IND ALTM VIB.
AC power is furnished by the 115 volt ac essential
bus. The altimeter encoder is protected by the
ALT ENCDR circuit breaker and the transponder
computer is protected by the IFF CMPTR circuit
breaker. To ensure reliable operation, a one

minute stabilization period is recommended with
the rotary master switch in STBY position prior
to system operation.

TRANSPONDER CONTROL PANEL.

The transponder control panel C-6280(P)/APX
(figure 7-8) is located in the pilot’s right console
and contains all the operating controls with the
exception of the IFF code hold switch. The five
position rotary master selector switch controls
operation of the system as follows:

OFF — Identification system is de-energized.

STBY — Full power supplied to the system, but
interrogation is blocked.

LOW — Receiver sensitivity is reduced by preset
amount, permitting only high energy signals to
trigger the transponder.

7-15

Section VII

NAVAIR 01 -HIAAB-1

Figure 7-8. Identification Transponder (Sheet 1 of 3)

7-16

NAVAIR 01 -HIAAB-1

Section VII

NOMENCLATURE

1. MASTER Control OFF

STBY

LOW

NORM

EMER

FUNCTION

Turns set off.

Places in warmup (standby) condition.

Set operates at reduced receiver sensitivity.

Set operates at normal receiver sensitivity.

Transmits emergency reply signals to MODE 1,

2, or 3/A interrogations regardless of mode control
settings.

RAD TEST - MON Switch

RAD TEST

Enables set to reply to TEST mode interrogations .

Other functions of this switch position are classified.

MON

Enables the monitor test circuits.

OUT

Disables the RAD TEST and MON features.

IDENT-MIC Switch

IDENT

Initiates identification reply for approximately 25 seconds

OUT

Prevents triggering of identification reply.

Spring loaded to OUT.

MIC

Initiates identification reply for approximately 25
seconds when the AN/ARC-159 transmitter is keyed.

MODE 3/A Code

Select Switches

Selects and indicates the MODE 3/A four-digit
reply code number.

MODE 1 Code Select
Switches

Selects and indicates the MODE 1 two-digit reply
code number.

MODE 4 Switch

Enables the set to reply to MODE 4 interrogations.

OUT

Disables the reply to MODE 4 interrogations.

AUDIO-LIGHT Switch

AUDIO

Enables aural and REPLY light monitoring of valid

MODE 4 interrogations and replies.

LIGHT

Enables REPLY light only monitoring of valid MODE 4
interrogations and replies.

OUT

Disables aural and REPLY light monitoring of valid

MODE 4 interrogations and replies.

CODE Control

Functions of this switch are operationally classified.

M-1 Switch

Enables the set to reply to MODE 1 interrogations.

OUT

Disables the reply to MODE 1 interrogations.

TEST

Provides test of MODE 1 interrogations by indication
on TEST light.

210077-64-2B

Figure 7-8. Identification Transponder (Sheet 2 of 3)

7-17

Section VII

NAVAIR 01 -HI AAB-1

NOMENCLATURE

10.

REPLY Indicator

11.

M-2 Switch

OUT

TEST

12.

TEST Indicator

13. M-3 A Switch ON

OUT

TEST

14. M-C Switch ON

OUT

TEST

FUNCTION

Lights when valid MODE 4 replies are present, or
when pressed.

Enables the set to reply to MODE 2 interrogations.

Disables the reply to MODE 2 interrogations.

Provides test of MODE 2 interrogations by indication
on TEST light.

Lights when the set responds properly to a M-1, M-2.
M-3/A or M-C test, or when pressed.

NOTE

Kit 1A/TSEC (classified) computer, must
be installed before set will reply to a MODE 4
interrogation.

Enables the set to reply to MODE 3/A interrogations.

Disables the reply to MODE 3/A interrogations.

Provides test of MODE 3/A interrogation by indication
on TEST light.

Enables altitude reporting in conjunction with AAU 32A.
Disables altitude reporting function.

Provides test feature for mode C function.

210077-64-3R

Figure 7-8. Identification Transponder (Sheet 3 of 3)

NORM — Transponder will operate at normal
sensitivity and respond to interrogations in
accordance with settings of other controls.

EMER — Allows the system to respond to
interrogations in modes 1,2, and 3/A regardless of
the settings of the mode control toggle switches.
Modes 1 and 2 will respond with codes selected on
the applicable dials, plus a recognizable emer¬
gency pulse train. Mode 3/A will transmit code
7700 regardless of the code selected on the dial.
Modes C and 4 will respond normally, regardless
of the position of the selector switches. A detent
prevents accidental selection of the EMER position
and is bypassed by raising the selector knob.

The four mode control toggle switches marked
M-1, M-2, M-3/A and M-C have three marked
positions, OUT, ON, TEST. The OUT (down)

position prevents responses and the ON (center)
position permits responses in the mode selected.
The spring loaded TEST (up) position provides a
self-test capability in conjunction with the on board
transponder self-test set. When one of the mode
control toggle switches is held momentarily (2-3
sec) in the TEST position, the on board self-test
set will generate interrogation pulses pairs for the
mode selected. These interrogations are applied to
the transponder to check for proper receiver
frequency, sensitivity, and decoding. A reply from
the transponder indicates that the transponder is
tuned to the correct receiver frequency and has
normal sensitivity. The test set analyzes the
replies to ensure that spacing of bracket pulses,
transmitter frequency, power, and antenna circuit
VSWR (threshold for rejection) are all above the
preset minimum acceptable standard. If the
characteristics of the reply transmission are

7-18

NAVAIR 01-H1AAB-1

Section VII

within the preset limits, the test set will illuminate
the TEST light on the transponder set control
panel, indicating to the pilot that the AIMS
Transponder System is functioning normally and
a “go” condition exists. If any one or more of the
transmission characteristics is not within pre¬
scribed limits, the TEST lamp will not illuminate
indicating that a “NO GO” (improper operation)
condition exists. The rotary master switch must be
set to NORM for the test function to operate. The
mode switches of the modes not being tested
should be OUT when testing on the ground to
prevent transponder operation on the non-tested
modes. The TEST light may flash once as each
mode switch is released from the TEST position,
and as the RAD TEST-OUT-MON switch is moved.
This is a characteristic of the transponder self test
set (TS-1843/APX) and is meaningless.

The mode 1 and 3/A code selector dials are small
rotatable drums with imprinted numbers which
appear through the code selector windows. The
numbers are changed by rotating the drums by
means of the raised tabs. Mode 1 provides 32
possible combinations from 00 to 73. Mode 3/A
provides 4096 possible code combinations from
0000 to 7777. Mode 2 code settings are preset on
the transponder panel located in the tail boom of
the aircraft. Mode 2 provides 4096 possible code
combinations from 0000 to 7777.

The IDENT-OUT-MIC switch is a three position
toggle switch located at the lower right corner of
the transponder control panel. The spring loaded
IDENT position provides an identification reply
on selected modes for 15 to 30 seconds after
releasing the IDENT switch. When MIC position
is selected, the identification reply activation is
transferred from the IDENT switch to the UHF
microphone switch causing an identification reply
each time the UHF microphone switch is depressed.
The MIC position is not spring loaded and must be
manually repositioned to OUT. The OUT position
de-energizes the identification reply circuit.

The three position RAD TEST-OUT-MON toggle
switch is located on the right center of the trans¬
ponder control panel. The RADiated TEST
position enables ground avionics maintenance
personnel to test mode 3/A transponder replies
when using test set interrogations from an AN/
UPM-92 or similar ground test equipment. Other
functions of this position are used with mode
4 test. In MON position, the switch energizes
monitor circuits of the on board transponder
self-test set. The TEST light will illuminate for 3

seconds each time an acceptable reply is trans¬
mitted in response to normal external interroga¬
tions in any selected mode. The OUT position
de-energizes both the RAD TEST and MON
circuits.

IFF CONTROL PANEL.

The IFF control panel located to the right
of the pilot, contains the switches and controls
for applying power and selecting the mode
of operation of the transponder equipment
figure 7-8.

Transponder Operating Procedure.

1. IFF XPONDER, IFF TEST and IFF CMPTR
circuit breakers — CLOSED.

2. MASTER control — STBY (for 3 minutes) —
LOW/NORM.

3. M-l, M-2, M-3/A, M-C and MODE 4 - ON -
(unless operational requirements indicate
that only specific modes will be used, then all
other switches will be OUT.)

4. AUDIO-LIGHT - LIGHT.

5. IDENT-MIC - OUT.

6. RAD TEST-MON - OUT.

7. INTER IFF — ON (only when mode 4 equip¬
ment is installed and operating).

8. To secure set MASTER OFF.

Radar Altimeter - AN/APN-17KV).

The AN/APN-171(V) radar altimeter is an airborne,
high resolution, short pulse, tracking and indicating
radar system. The set measures and visually
indicates actual clearance in feet between the
helicopter and terrain, over a range from 0 to 5000
feet. The set consists of a receiver-transmitter
installed in the tailboom, a receive antenna and a
transmit antenna both located on the lower surface
of the tailboom and a height indicator located in
the pilot instrument panel. The low altitude
warning will be dimmed for night operation using
the caution lamp dimming switch if pilot instru¬
ment lights are on.

7-19

Section VII

NAVAIR 01 -HI AAB-1

AN/APN-17KV) CONTROLS.

All control of the AN/APN-171(V) altimeter is
centered in a single control knob located on the
indicator (figure 7-9). Control functions are:

1. Low altitude warning position is selected by
rotation of the low altitude set knob and
indicated by a pointer traveling the outer
circumference of the indicator dial.

2. System power is turned off by rotating the low
altitude set knob fully counterclockwise so that
the index is below zero altitude.

3. Depressing the low altitude set knob activates
the system’s self test feature.

Radar Altimeter Operating Procedure.

1. RADAR ALTM and RADAR ALT circuit
breakers — CLOSED.

2. PUSH TO TEST knob — Rotate clockwise
from off to desired LOW altitude caution
light setting.

3. Test system by depressing low altitude set knob.
When knob is depressed, a reading of 100 plus or
minus 15 feet will be indicated if the system is
functioning properly. Releasing the push to test
knob restores the system to normal operation.

Bearing-Distance-Heading Indicator (BDHI).

There are two bearing-distance-heading-indicators
(BDHI) mounted in the helicopter; one on the
pilot instrument panel (figure 7-10) and one on
copilot/gunner instrument panel. They are connected
in parallel as repeaters and display the following
information:

Compass Card — provides an accurate indication
of the helicopter heading and is controlled by the
AN/ASN-75 gyro compass system. Pointer No. 1
(single bar) displays either LF-ADF bearing from
the AN/ARN-83 system or UHF-DF bearing from
the AN/ARA-50. The AN/ARN-83 signal is
removed from the pointer when the ADF position
of the AN/ARC-159 is selected, UHF-DF bearing
will be displayed only.

Pointer No. 2 (double bar) displays the TACAN
bearing from the AN/ARN-84(V) system.

Digit numbers window displays the distance in
nautical miles to the selected TACAN station.

AN/ARN-84 (V) TACAN.

The TACAN set is an air navigation system (figure
7-11). TACAN provides range and magnetic bearing
to the selected station. The set transmits and receives
radio frequency signals, demodulates and decodes
the received signals, computes bearing and slant
range within 300 nautical miles, processes beacon
identification signals, and processes the self test
signals when a self test command is initiated at the
control panel.

TACAN receives on 126 channels with X or Y coding.

The bearing signal is displayed on the pilot and
copilot/gunner BDHI No. 2 pointer. Line-of-sight
distance range is displayed in nautical miles on the
distance range indicator of each respective BDHI.
The course deviation R-L and deviation warning
signals are presented to the pilot on the ADI. When
the VHF-FM is not in the homing mode, TACAN
course deviation and flag signals will be presented
on the respective pointer and flag.

AN/ARN-118 (V) TACAN.

The TACAN set is an air navigation system (figure
7-8A) which operates by transmitting and receiving
radio frequency signals. The received signals are
demodulated and decoded to compute slant range
distance and magnetic bearing to/from the selected
station. The bearing information is displayed by
the No. 2 pointer on the pilot/copilot BDHI, while
the line-of-sight distance range is displayed in
nautical miles on the distance range indicator of
each respective BDHI. The course deviation R-L
and distance warning signals are presented to the
pilot on the ADI. When the VHF-FM is not in the
homing mode, TACAN course deviation and flag
signals will be presented on the respective pointer
and flag.

TACAN receives on 126 channels with X or Y
coding.

The TCN-118 TACAN set contains an automatic
self-test function that causes the system to be
tested automatically when the TACAN beacon
signal is lost. The automatic self-test checks the
system for proper operation to determine if the
signal loss was due to a system malfunction. If
there is a system malfunction, the TEST indicator
on the TACAN control panel illuminates upon
completion of the automatic self-test cycle.

7-20

Change 1

NAVAIR 01-H1AAB-1

Section VIII

To provide an inflight confidence test, a manual
self test can be initiated by momentarily depressing
the TEST switch on the control panel. When
initiating a system test, observe the control TEST
indictor. If the indicator illuminated during the
test and remains illuminated, there is a malfunction
in the system and the information displayed on the
BDHI should be disregarded. If the TEST light
extinguishes, the system is checked and provides an
85 percent confidence level.

One other advantage of the TCN-118 over previous
TACAN systems is that it provides both bearing
and distance information in the air-to-air mode
between aircraft operating with compatible systems.

The following represents operating range over
level terrain:

AGL (FEET)

DME

100

250

500

750

1000

2000

3000

4000

5000

TACAN Set — Operating Procedures.

1. ICS PLT and GUNNER circuit breakers —
CLOSED.

2. CMPS and ADF IND circuit breakers —
CLOSED.

3. TACAN dc, TACAN SYS ac, and TACAN
XCVR circuit breakers - CLOSED.

4. GYRO CMPS circuit breaker — CLOSED.

5. INV-MAIN and STBY circuit breakers —
CLOSED.

6. BATTERY - ON.

7. INVERTER - MAIN.

8. ICS control panel TACAN monitor — ON
(OUT).

9. TACAN MODE selector - AS REQUIRED.

10. TACAN CHAN selectors-AS REQUIRED.

NOTE

Depressing the “BITE” switch on the
TACAN control panel will interrogate the
system with a test signal and provide
indication of the operational condition by
means of the GO-NO GO status lamps. A
GO lamp indicates the system to be
operational and will remain illuminated
for about nine seconds. A NO GO lamp
indicates that the system is not operating
correctly and maintenance action is
required. The NO GO lamp will remain
illuminated as long as power is applied to
the system.

Change 1

7-20A

Section VII

NAVAIR 01-H1AAB-1

CHANNEL DIGITAL VOL

DISPLAY CONTROL

CHANNEL SELECTOR CONTROLS

NOMENCLATURE

FUNCTION

OFF

Off switch for TCN-118 system.

REC

Receive Mode: Provides relative bearing.

T/R

Transmit-Receive Mode: Provides bearing and distance information.

A/A REC Air to air receive mode: receives bearing information from

suitably equipped, cooperating aircraft (no distance supplied).

A/A T/R Air to air transmit-receive mode: receives both bearing and

distance information from a suitably equipped, cooperating aircraft.
If the cooperating aircraft is not equipped with bearing trans¬
mitting capability, only slant range distance to the aircraft
is provided.

CHANNEL selectors Selects any of 126 channels.

X-Y

Selects either X or Y channel.

UNITS selector
TENS/HUNDREDS
CHANNEL digital display
VOL control
TEST switch
TEST indicator

Selects units digit of desired channel.

Selects tens and/or hundreds digits of desired channel.

Displays TACAN channel.

Varies level of audio signal.

Indicates system self-test or confidence test (pilot induced).

Illuminates when a malfunction occurs during either test (flashes
at start of self-test cycle to check indicator lamp).

Figure 7-8A. ARN-118 (U) TACAN

7-20B

Change 1

NAVAIR 01 -HI AAB-1

Section VII

NOMENCLATURE

Circuit breakers
Low altitude bug
PUSH TO TEST

POWER OFF
NO TRACK FLAG

NO TRACK MASK

LOW ALTITUDE CAUTION

FUNCTION

Protects individual circuit.

Indicates selected altitude.

PUSH TO TEST — Test altimeter for operation.

SET — Selects altitude at which low altitude warning light
operates.

OFF — Removes power. Indicates power removed from
altimeter or malfunction.

Indicates altimeter unreliable, or above 5000 feet AGL.
Illuminates when below selected altitude.

210077-69B

Figure 7-9. Radar Altimeter

7-21

Section VII

NAVAIR 01-H1AAB-1

COURSE SELECTOR
INDICATOR

POWER OFF
FLAG

“POINTER NO. 2
(DOUBLE BAR)

DISTANCE RANGE
INDICATOR

COURSE (CRS) CONTROL
KNOB

NOMENCLATURE
Power OFF Flag.

Course Selector Index (BUG)
Fixed Index

Pointer No. 1 (single bar)
Pointer No. 2 (double bar)
Distance Range Indicator

Course (CRS) control knob
positions the bug

FUNCTION

Disappears when gyro.nagnetic compass is energized

Reference for course deviation pointer on ADI.

Gyromagnetic compass heading on compass card
card under lubber line.

Indicates ADF/UHF/DF bearing

Indicates TACAN bearing.

Line-of-sight distance (nautical miles) to TACAN
station

Positions the bug and selects course selected by the
ADI deviation pointer.

21090 0-70A

Figure 7-10. Bearing-Distance-Heading Indicator

7-22

NAVAIR 01 -HIAAB-1

Section VII

BITE SELF TEST MODE

FUNCTION

Illuminate for nine seconds, upon completion of no-fault self-test.
Illuminate, for indefinite period, when a fault occurs during selt-test.

Initiate self-test when depressed and released.

Not used.

Selects audio level of the beacon identity tone.

Receives beacon signal and computes bearing between helicopter and
surface beacon.

Receives beacon signal and computes slant range and bearing
between helicopter and surface beacon.

Computes slant range between two aircraft equipped with
TACAN. Shall be 63 channels apart.

Becomes an omnidirectional beacon for other aircraft. (Non-functional)

Selects any one of 126 TACAN transmit channels.

210900-146

Figure 7-11. Tacan

7-23

NOMENCLATURE

STATUS lights
GO
NO-GO

BITE switch

MODE control switches

VOL control

Mode selector switch
REC

T/R

A/A

BCN

CHAN selector switches

Section VII

NAVAIR 01 -HIAAB-1

a. Identify tone output present for
approximately 10 seconds.

b. The No. 2 bearing pointers shall indicate
between 0° and 8° for approximately 15
seconds, followed by search.

c. Range between 1.6 nm and 2.0 nm for
approximately 20 seconds, followed by
search.

d. The GO lamp shall light after
approximately 21 seconds, and remain so
for approximately 9 seconds.

11. TACAN BITE - DEPRESS, RELEASE.

12. TACAN AND ICS volume controls — AS
REQUIRED.

AN/ARN-83 Automatic Direction Finder.

The direction finder set provides radio aid to
navigation. It operates in the frequency range of
190 to 1750 KHz. When operating as an automatic
direction finder, the system presents a continuous
indication of the bearing to any selected radio
station. It also provides simultaneous aural
reception from the station. When the manual
(LOOP) mode of operation is selected, the system
enables the operator to find the bearing to any
selected radio station by manually controlling the
null direction of the directional antenna. The
system also operates as a radio range receiver and
a conventional low-frequency aural receiver to
receive voice and unmodulated transmission.
Information received via the direction finder set is
presented visually on the No. 1 needle of the pilot
and copilot/gunner BDH indicators, and aurally
through the intercom system. The receiver is
installed in the aft radio compartment. Electrical
power to operate the set is supplied by the 26 volt ac
essential bus through circuit breakers marked ADF
RCVR and ADF IND from the 28 volt dc non-
essential bus through a circuit breaker marked
ADF RCVR.

ADF CONTROL PANEL.

The ADF control panel is marked ADF and is located
in the pilot right console (figure 7-12). The panel
incorporates the controls for the ADF receiver and
associated loop antenna and sense antenna.

ADF Operating Procedure.

To operate the direction finder set in any particular
mode, perform the following preliminary steps:

1. ADF RCVR, ICS PLT, INV MAIN dc circuit
breakers — CLOSED.

2. ADF RCVR, ADF IND, GYRO CMPS,
28 vac circuit breakers — CLOSED.

3. ADF monitor — PULL OUT, ADJUST.

4. Mode selector — AS DESIRED.

5. Frequency — SELECT.

ADF OPERATION.

1. Mode selector — ADF.

2. BFO-OFF — OFF (Except for CW station).

3. Tuning meter — TUNE (for maximum
deflection.)

4. GAIN control — ADJUST.

ANTENNA OPERATION.

In this mode the No. 1 pointer on the BDH indicator is
inoperative. To operate the ADF set in the ANT mode
perform the following:

1. Mode selector — ANT.

2. GAIN control — ADJUST.

MANUAL LOOP OPERATION.

1. Mode selector — LOOP.

2. BFO-OFF — OFF (Except for CW station).

3. GAIN control — ADJUST.

4. LOOP — MOVE (left or right for null.)

7-24 Change 1

NAVAIR 01 -HI AAB-1

Section VII

NOMENCLATURE FUNCTION

Band selector
switch

TUNE control
Tuning meter

GAIN control

Selects desired frequency band. BFO switch

Turns BFO on or off.

Mode

Selects the desired frequency. selector

switch

Facilitates accurate tuning of
the receiver.

Controls receiver audio volume.

ADF — Automatic direction finding
station direction on No. 1
pointer in BDHI.

ANT — Low frequency radio
receiver using sense
antenna only.

LOOP L-R Rotates loop antenna to the

right or left when in loop
mode.

LOOP — Manual direction finding
or aural null operation using

loop antenna only. ,

210900-145

Figure 7-12. Automatic Direction Finder (ADF)

7-25/(7-26 blank)

NAVAIR 01 -HI AAB-1

Section VIII

SECTION VIII — WEAPONS SYSTEM

TABLE OF CONTENTS

Introduction.8-1

Armament Configuration.8-1

Interrelation of Armament.8-1

Armament Firing Modes.8-1

GTK4A/A Turret System .8-3

Pilot Armament Controls and

Indicators .8-5

Gunner Armament Controls and

Indicators .8-11

Cyclic Stick Armament Switches.8-11

Helmet Sight Subsystem (HSS)....8-17

TOW Missile System (TMS).8-19

TSU Guns.8-30

Wing Stores Jettison.8-32

Wing Stores Armament System.8-32

Rockets.8-32

Wing Gun Pod.8-32

Smoke Grenade Dispenser.8-32

Preflight Procedures.8-32

Inflight Procedures — All Armament.8-37

Turret Operation.8-37

Post Firing/Before Landing Check

— All Armament. 8-40

Countermeasures Dispensing System

(AN/ALE-39).8-40

Radar Warning System (AN/APR-39).8-40

Countermeasures System AN/ALQ-144.8-41

NOTE

• In this section, the copilot/gunner will
be referred to as gunner for brevity.

• Refer to figure 8-22 for acronyms used
in this section.

INTRODUCTION.

The AH-1T (TOW) helicopter provides a high
degree of armament carrying versatility through
the utilization of six integral armament control
systems: GTK4A/A Turret System, Navy

Armament Control and Delivery System, Emer¬
gency Jettison System, Smoke Grenade Dispenser
System, TOW Missile System and Helmet Sight
Subsystem.

• A single store may be flown within the
6-inch lateral CG limits. Stores on both
store stations on the same side, with
opposite side empty, can exceed the
lateral CG limit for some loadings.

• Firing of TOW missile, 2.75-inch FFAR,
and 20-mm gun in icing conditions is
prohibited. The TOW missile warhead
can detonate in close proximity to the
helicopter. The warhead fuse can be
damaged as missile is launched through

ice in missile launcher. Gun barrels and
breeches can rupture if gun muzzles are
clogged with ice. FFAR can be held
captive in the launcher tubes by the
frozen ice.

• Helicopter control shall be maintained,
especially at low altitude to prevent
hazardous flight conditions and loss of
TOW missile control. When gunner is
tracking a TOW missile and pilot is
using his helmet sight to fire the turret
simultaneously, the pilot may have a
strong tendency to lose contact with his
instrument panel and outside references
or develop target fixation.

ARMAMENT CONFIGURATION.

Refer to the AH-1 TACTICAL MANUAL, NAV¬
AIR 01-110HC-1T, for all the authorized store
loadings.

INTERRELATION OF ARMAMENT.

The armament subsystems are interfaced with
one another. Figure 8-1 shows pilot and gunner
control components in relationship to each arma¬
ment subsystem.

ARMAMENT FIRING MODES.

Figure FO-8 shows switch positions for principal
firing modes.

Change 1

8-1

Section VIII

NAVAIR 01-H1AAB-1

CONTROL

COMPONENTS

TURRET

TOW

MISSILE

WING STORES
GUN

ROCKETS POD

SMOKE

GRENADE

DISPENSER

TARGET
ACQUIRE
FOR TSU

—->

WING

STORES

JETTISON

PILOT STATION:

Armament Control Panel

Store Control Panel

Smoke Grenade Dispenser
Control Panel

Smoke Grenade Release
Switch

Wing Stores Jettison

Button

Pilot Steering Indicator

Gunner Accuracy Control
Panel

Training

Fixed Sight

Helmet Sight

Cyclic Switches

Emergency Jettison

Select Panel

GUNNER STATION:

Cyclic Switches

Helmet Sight

Telescopic Sight Unit

Left Hand Grip

Armament Control Panel

Wing Stores Jettison

Switch

Sight Hand Control

TOW Control Panel

Figure 8-1. Interrelation of Armament

8-2

NAVAIR 01-H1AAB-1

Section VIII

GTK4A/A TURRET SYSTEM.

The GTK4A/A Turret System provides for
positioning, sighting ammunition feeding, and
firing of the M197 20-mm automatic gun. The
system consists of turret assembly, turret control
assembly, pilot and gunner controls, pilot and
gunner helmet sight, gunner TSU, airspeed pres¬
sure transducer, ammunition feed system with
booster assembly, gun control assembly, gun drive
assembly and a gun recoil adapter. If electrical
power is removed from the Turret System for
any reason, emergency 28 vdc power is applied
directly to an auxiliary drive circuit in the turret
assembly which automatically brings gun to upper
stow position for safe landing. Upper stow position
for gun is zero degrees azimuth, 11 to 14 degrees
up elevation, and lower stow position is zero
degrees azimuth and elevation. A pilot override
mode is provided, which allows gunner emergency
operation of the weapons system, less smoke
and TOW.

Functions.

The turret assembly, which is chin-mounted on the
helicopter (figure 8-2), provides mounting for the
M197 20-mm automatic gun. Electrical circuits in
the turret control assembly and turret assembly
provide remote control for azimuth and elevation
drive system in the turret. The azimuth drive
system rotates the turret through a range of 110
degrees either side of zero degrees azimuth. Gun
can be lowered 50 degrees below zero degrees
elevation. The gun control assembly controls
operation of gun and operation of the ammunition
system. The gun control assembly also supplies
firing voltage and supplies it to gun when gun
drive power is applied. The gun drive assembly
rotates gun barrels at a firing rate of approxi¬
mately 650 rounds per minute. Power applied
to the ammunition feed system operates booster
assembly on ammunition box and energizes a
declutching solenoid on gun feeder. The booster
assembly and gun feeder provide a flow of
ammunition from the ammunition box. The
ammunition feed system contains 750 rounds of
belted 20-mm ammunition. The gun is fired for
duration of trigger command signal plus clearing
cycle, or in limited 16 ± 4 round bursts. The
first detent on the cyclic trigger fire switch/LHG
trigger allows gun control assembly to auto¬
matically, terminate each trigger command signal
after 16 ± 4 rounds are fired. The second detent
allows continuous firing.

CAUTION

The M197 automatic gun is restricted to
a firing schedule not to exceed a 450
round burst with a minimum of 6
minutes cooling time prior to firing
remaining 300 rounds.

The gun control assembly also terminates trigger
command signal when gun reaches azimuth or
elevation limits, and when gun position disagrees
in azimuth or elevation more than 5.5 degrees
from sight position command signal. Flow of
ammunition to gun stops immediately upon
termination of trigger command signal. Time delay
circuits in gun control assembly continue gun drive
power and firing voltage long enough for gun to
fire ammunition remaining in gun unless trigger
command signal is terminated by position error.
Should trigger command signal be terminated by
position error in excess of 5.5 degrees, firing
voltage is terminated simultaneously; however,
gun drive power is continued to clear live ammuni¬
tion from gun.

The turret has five modes of operation. Pilot and
gunner control utilizing the HSS are fully covered
under HSS in this section. The TSU guns mode is
fully covered under TMS in this section. The pilot
or gunner may fire the turret in the fixed (Lower
Stow) position utilizing the respective cyclic trigger
fire switches when weapon control switch is in
fixed or pilot override is selected.

The turret and wing stores cannot be fired
simultaneously. Fixed forward mode: If the turret
is firing and the WING ARM FIRE button is
pressed, the clearing mode of the turret is actuated.
Gunner-mode: If the turret is firing and the pilot
TRIGGER ACTION switch is depressed, the
clearing mode of the turret is activated and gun
will stow. Rocket firing circuits are energized
allowing pilot to fire rockets using WING ARM
FIRE button.

The turret may continue firing approxi¬
mately 1/2 seconds after the WING ARM
FIRE button and pilot TRIGGER
ACTION switch have been released.

8-3

Section VIII

NAVAIR 01 -HI AAB-1

209071-391B

Figure 8-2. Turret

8-4

NAVAIR 01 -HIAAB-1

Section VIII

Pilot Mode.

If the turret is firing, pilot TRIGGER ACTION
switch must be released in order to fire rockets.
The pilot COMP switch provides recoil compensa¬
tion inputs to tail rotor and cyclic through the
SCAS. RANGE KNOBS on the pilot instrument
panel and gunner armament control panel provide
turret gun elevation inputs to compensate for
range selected.

PILOT ARMAMENT
INDICATORS.

CONTROLS AND

Refer to figure 8-3.

Pilot Armament Control Panel.

The panel contains the controls and indicators to
arm and fire armament subsystems and use the
helmet sight subsystem. Refer to figure 8-4.

MASTER ARM SWITCH.

The MASTER ARM switch is a three position
switch which permits the pilot to energize and
deenergize the armament circuits. Placing the
switch in ARM, arms the armament system. Placing
the switch in STBY, energizes the control circuits
for complete operation of weapons systems with
the exception of the trigger fire circuits. Placing
the switch in OFF, deenergizes the armament and
control circuits.

Smoke grenades can be released, utilizing the
smoke release button while master arm is in STBY
or ARM.

WEAPON CONTROL SWITCH.

A three position WEAPON CONT switch is
provided to allow the pilot to select the mode of
operation of the armament system. In the gunner
position, the gunner is primary armament system
operator.

The weapon CONT switch must be in the gunner
position to operate the TMS. However, the pilot
can also select and fire wing stores utilizing the
rocket lock-out (Pilots’ Trigger Turret Action). If
the gunner places the TCP mode select in any of

the TOW functions, control of the turret reverts
to the Pilots’ Helmet sight (PHS). In the FIXED
position, the pilot is in control of the turret and
wing stores. This position is used in conjunction
with the pilot fixed sight. In PILOT position, the
pilot is the primary armament system operator,
and has the capability to fire the turret with the
HSS, and select and fire wing stores. For a
depiction of armament firing modes, refer to FO-8.

Navy Armament Control and Delivery System
(NARCADS).

The NARCADS stores control panel (figure 8-5)
enables the pilot to program the automatic release
of weapons/stores from four wing stations in the
quantity, mode, and rate selected; and the gunner
to select, in PILOT OVERRIDE condition, inboard
or outboard stations only. The system permits
selective weapons/stores release by the pilot. It
incorporates safety features which require
matching the selected station and the selected
type of weapons/stores registered for that station.
NARCADS provides capability for inflight arming
of droppable weapons.

NOTE

• Incorrect thumbwheel settings may
allow weapons to fire.

• Identical weapons may be released/
fired simultaneously from wing stations.
There is no capability to fire dissimilar
weapons simultaneously.

• By selecting a dissimilar store on the
station select, the previously selected
dissimilar store will automatically
deselect. The pilot can select additional
stations for the gunner. Example: If the
gunner selects Pilot override, inboard,
the pilot can additionally select stations
1 and/or 4. If the QTY counter shows
0/0 in SINGLE or PAIR mode, no
weapons will fire.

Pilot Fixed Sight.

Refer to figure 8-6.

8-5

Section VIII

NAVAIR 01-H1AAB-1

N2/83

210071-63

Figure 8-3. Pilot Controls and Indicators

8-6 Change 1

NAVAIR 01 -HIAAB-1

Section VIII

LOCATION: PILOT INSTRUMENT PANEL

NOMENCLATURE

MASTER ARM Switch OFF

STBV

ARM

WEAPON CONT Switch FIXED

PILOT

GUNNER

ARMED STBY Indicator ARMED

STBY

Off

Press

FUNCTION

Deactivates all sights and weapon control/firing circuits.

Activates all sights, turret and TOW missile control circuits, and arms smoke
grenade firing circuits. Charges wing gun pod battery. Illuminates pilot arid
gunner STBY lights.

Activates all sights and weapon control/firing circuits. Chargeswing gun pod
battery. Illuminates pilot and gunner ARMED lights.

Permits pilot to fire turret and wing stores (not TOW) using fixed sight.
Permits pilot to fire turret using HS and wing stores (notTOW) using fixed sight.
Permits gunner to fire turret using HS or TSU and TOW using TSU. Illuminates
GUNNER IN CONT light on gunner armament panel.

Indicates MASTER ARM switch in ARM (amber light) or gunner PILOT
OVERRIDE switch is in OVERRIDE.

Indicates MASTER ARM switch in STBY (green light).

Indicates MASTER ARM switch is off.

Tests indicator lights.

.210075-209

Figure 8-4. Pilot Armament Control Panel

8-7

Section VIII

NAVAIR 01 -HI AAB-1

NOMENCLATURE

FUNCTION

THUMBWHEEL REGISTRATION
SWITCHES

NOTE

Switches must be manually set to indicate
weapons installed before weapons can be
fired.

EMPTY

— Indicates no weapons (Off)

RKT 7

— 7-tube rocket launcher

RKT 19

— 19-tube rocket launcher

GUN PD

- GPU-2A

HTW

— Not used

FLARE

— Flare

TNG BB

— Training bomb

BOMBS

— Bomb

RKT 4

— 4-tube rocket launcher

DISP

— SUU-44 dispenser

(blank)

— Not used

TOW

— Indicator only for TOW rack

210071-31-1

Figure 8-5. Pilot Store Control Panel (Sheet 1 of 2)

8-8

NAVAIR 01 -HI AAB-1

Section VIII

NOMENCLATURE

FUNCTION

STATION SELECT

SELECT

— Indicates wing station is selected for firing and weapons are present.

— (For Rockets Only) . ,

Indicates five or less rockets remain in 7-tube or 19-tube launcher.

- Indicates weapons are depleted, except for gun pod and TOW.

CP CP

- Indicates that PILOT OVERRIDE mode is selected and gunner controlsturret and
wing stores except TOW. Previous station selections are deselected. RATE, UI Y,
MODE, and thumb wheel settings remain applicable.

BOMB ARM SAFE

— All arming solenoids de-energized.

TAIL

— Tail arming solenoids energized. Nose arming solenoids de-energized.

NOSE

— Nose arming solenoids energized. Tail arming solenoids de-energized.

BOTH

— Both nose and tail arming solenoids energized.

RATE. FAST

— (For Rockets Only)

Sets release rate at 90 milliseconds.

SLOW

- Sets release rate at 180 milliseconds.

QTY

THUMBWHEELS/

INDICATOR

mode SINGLE

— Selects and displays number of fire pulses to be generated in the SINGLE or PAIR
mode. (This function is disabled on ALL mode).

— Opposite wing station weapons of selected like stores to fire alternately.

PAIR

— Opposite wing station weapons of selected like stores to fire concurrently.

ALL

- All selected wing station weapons to fire concurrently. QTY select function
disabled.

210071 - 31-2

Figure 8-5. Pilot Store Control Panel (Sheet 2 of 2)

Section VIII

NAVAIR 01 -HI AAB-1

Figure 8-6. Pilot Sight

8-10

Change 1

NAVAIR 01-H1AAB-1

Section VIII

Pilot Armament Circuit Breakers.

Refer to figure 8-7.

Smoke Grenade Dispenser Control Panel.

The SMOKE control panel is located on the pilot’s
left bulkhead (figure 8-8). The color of smoke
grenades in each rack is set on a color indicating
dial located below the ARM switch for each rack.
To select grenades of a desired color, the pilot
actuates the ARM switch directly above the color
indicating dial set to the desired color. Pressing
the SMOKE REL button on the pilot collective
switch box, drops one grenade from each of four
racks selected, and initiates a 400-cycle audio
tone in the pilot’s headset. The pilot hears the
same audio tone regardless of how many grenades
he is firing simultaneously and the tone is heard as
long as the SMOKE REL button remains depressed.
When the last grenade from the rack had been
fired, the tone continues until the ARM switch
for the particular rack is placed OFF.

GUNNER ARMAMENT CONTROLS AND
INDICATORS.

Refer to figure 8-9.

Gunner Armament Control Panel.

The panel (figure 8-10), located on the gunner’s
right console, contains controls and indicators
which enable the gunner to operate and monitor
armament subsystems. The gunner can take
armament command, regardless of pilot MASTER
ARM switch position, through the use of the
PILOT OVERRIDE switch. The gunner can then
fire the turret and wing stores (not TOW and
SMOKE) by use of his cyclic stick armament
switches.

PILOT OVERRIDE SWITCH.

The PILOT OVERRIDE switch is located on the
gunner armament control panel (figure 8-10).
When in OVERRIDE, the switch electrically
bypasses the MASTER ARM switch on the pilot
armament control panel. All pilot store control
panel switch positions remain valid. The CP light
on the pilot store control panel illuminates,
GUNNER IN CONT light illuminates, pilot and
gunner ARMED lights illuminate and TCP status
annunciator displays OFF. Armament systems are

armed and controlled by the gunner, with the
exception of TOW and smoke grenade release. The
TSU is disabled during PILOT OVERRIDE opera¬
tion; however, the turret can be controlled by the
HS. The gunner can then fire the turret by depres¬
sing both cyclic TRIGGERS. Wing stores can be
released or fired by placing gunner armament
control panel WING STORES SELECT switch
in INBD or OUTBD and depressing cyclic WING
ARM FIRE button. When gunner places WING
STORES SELECT switch to INBD, stations 2
and 3 are selected; when placed to OUTBD stations,
1 and 4 are selected. Placing the switch to OFF
deselects all previously selected stations. OFF
position of PILOT OVERRIDE switch allows
normal control and operation of armament
systems.

I^ARNINy

• When in PILOT OVERRIDE, the arma¬
ment system is armed and may be fired
regardless of position of MASTER ARM
switch.

• The gunner, while in pilot override,
may fire dissimilar stores if dissimilar
stores are loaded on opposite inboard
or outboard stations.

NOTE

Pilot can disable all armament circuits
by deenergizing the appropriate circuit
breakers.

CYCLIC STICK ARMAMENT SWITCHES.

The pilot and gunner cyclic stick provides three
armament switches, WING ARM FIRE, TRIGGER
TURRET FIRE and TRIGGER ACTION.

Wing Arm Fire.

The WING ARM FIRE button on the cyclic stick
is used to fire wing stores. After selecting wing
stores, wing stores may be fired. An interrupter
circuit, interrupts turret firing depending on
mode of operation, i.e., fixed, gunner, or pilot.

Change 1

8-11

Section VIII

NAVAIR 01 -HI AAB-1

DRIVE GUN EL
MOTOR MOTOR STOW

HSS SECU i—TUI

ooo

CONT CONT

AC/ARMAMENT CIRCUIT BREAKER PANEL
LOCATION: PILOT LEFT CONSOLE

CIRCUIT BREAKER FUNCTION - APPLIES POWER TO AND PROTECTS CIRCUIT FOR

BOMB ARM/PLT SIGHT
REF/XFMR
TMS/PWR
TURRET
DRIVE MOTOR
GUN MOTOR
EL STOW
LH WING
SMK GREN
INBD/GUN POD
OUTBD/GUN POD
RH WING
SMK GREN
INBD/GUN POD
OUTBD/GUN POD
HSS/PWR
SECU/PWR
TURRET
PWR
CONT
WEAPON
CONT
FIRE

WING STORES
PWR

JTSN/GNR

JTSN/PLT

Bomb arming and pilot sight reticle.

Reference power for weapon system.

TOW missile firing.

Turret control.

Weapon rotation
Turret stow.

Left hand smoke grenade dispenser.

Left hand inboard gun pod battery charging.
Left hand outboard gun pod battery charging.

Right hand smoke grenade dispenser.

Right hand inboard gun pod battery charging.
Right hand outboard gun pod battery charging.
Helmet sight subsystem.

Servo electronic control unit.

AC turret control.

DC turret control

Weapon control.

Weapon firing.

Weapons firing less smoke.

Gunner jettison.

Pilot jettison.

210075-210

Figure 8-7. Pilot Armament Circuit Breakers

8-12

NAVAIR 01 -HI AAB-1

Section VIII

LOCATION: PILOT LEFT BULKHEAD

NOMENCLATURE

FUNCTION

* LH ARM
Switches

* RH ARM
Switches

* Color
Indicators

OFF — Deactivates left wing smoke grenade circuit.

LH ARM — Permit pilot to fire left wing smoke grenades.

OFF — Deactivates right wing smoke grenade circuit.

RH ARM — Permit pilot to fire right wing smoke grenades.

B, Y, W, — Indicates color of grenades installed.

V, R,G

B — Blue

Y — Yellow
W — White

V — Violet
R - Red

G — Green

* One for each
rack of each
dispenser
(total of two
per dispenser)

210075-278

Figure 8-8. Pilot Smoke Grenade Dispenser Control Panel

8-13

Section VIII

NAVAIR 01 -HI AAB-1

CAMERA AND FILM
MAGAZINE PROVISIONS

210071-62

Figure 8-9. Gunner Armament Controls and Indicators

8-14

NAVAIR 01 -HI AAB-1

Section VIII

ARMED

STBY

PLT

GNR

EIA

GUNNER
IN CONT

TURRET

STOW

FUNCTION

— Indicates pilot MASTER ARM switch in ARM or
PILOT OVERRIDE in OVERRIDE.

— Indicates pilot MASTER ARM switch in STBY.

— Indicates failure in pilot HS

— Indicates failure in gunner HS

— Indicates failure in electronic interface assembly

— Indicates HSS operating properly

— Indicates PILOT WEAPON CONT switch in
GUNNER, or PILOT OVERRIDE in OVERRIDE.

— Indicates turret in upper stow position.

NOMENCLATURE

ROUNDS REMAINING
indicator
JETTISON
switch

SELECT

switch

HSS RETICLE
switches

FUNCTION

Displays number of rounds remaining for the turret weapon.
INBD — Selects inboard stores to be jettisoned.

BOTH — Selects all stores to be jettisoned.

OUTBD — Selects outboard stores to be jettisoned.

INBD — Selects inboard stores to be fired.

OFF — Deactivates gunner wing arm fire switch.

OUTBD — Selects outboard stores to be fired.

OFF-BRT — Adjusts light intensity of gunner HS reticle.

TEST — Tests gunner HS reticle lamps (3).

HSS BIT

PILOT OVERRIDE
switch

AIRSPEED COMP
switch

Tests HSS interface assembly and pilot and gunner linkage assembly
* — when both linkage arms are attached to BIT magnet.

OVERRIDE - Overrides pilot MASTER ARM switch. Permits gunner
to fire turret using HS and wing stores without sight
(Not TOW or smoke).

OFF — Permits pilot armament control panel to control weapons

OOMP — Applies airspeed data to turret positioning circuits.

OFF — Removes airspeed data from turret positioning circuits.

210071 - 30-1

Figure 8-10. Gunner Armament Control Panel (Sheet 1 of 2)

8-15

Section VIII

NAVAIR 01-H1AAB-1

NOMENCLATURE

FUNCTION

TURRET

— Limits downward travel to prevent turret weapons from striking ground

DEPR

LIMIT

switch

OFF

— Permits turret travel between minimum to maximum elevation.

TSU/GUNS

TRACK RATE
switch

HIGH

— Provides fast slew rates for TSU regardless of LHG MAG switch

position.

— Provides slow slew rates for TSU when LHG MAG switch in

LOW

LO position.

RANGE

switch

0-2000

— Provides meters — to—target data to compensation circuit.

210071 - 30-2

Figure 8-10. Gunner Armament Control Panel (Sheet 2 of 2)

8-16

NAVAIR 01-H1AAB-1

Section VIII

Trigger Turret Fire.

The TRIGGER TURRET FIRE switch on the
cyclic stick is used to fire the turret. After pre¬
setting switches on the pilot or gunner armament
control panels, the turret may be fired. The
TRIGGER ACTION switch on the cyclic stick
must be depressed prior to depressing the
TRIGGER TURRET FIRE switch. The gunner
| cyclic switches are energized only when his PILOT
OVERRIDE switch is in the OVERRIDE position.
A hinged guard prevents TRIGGER TURRET'
FIRE and TRIGGER ACTION switches from
being inadvertently depressed.

Trigger Action.

The TRIGGER ACTION switch on the cyclic stick
is used to slave turret movement to helmet sight
movement. A hinged guard prevents TRIGGER
ACTION and TRIGGER TURRET FIRE switches
from being inadvertently depressed. When gunner
is firing turret by use of LHG trigger, the pilot can
interrupt by depressing pilot TRIGGER ACTION
switch, placing WEAPON CONT switch on PILOT,
or FIXED, or placing MASTER ARM switch to
STBY or OFF.

HELMET SIGHT SUBSYSTEM (HSS).

The HSS (figure 8-11) permits pilot or gunner to
rapidly acquire visible targets and to direct turret
on to those targets.

The HSS also provides a means of cueing from
pilot to gunner for target location.

The HSS consists of two helmet sight (HS)
assemblies mounted on the pilot and gunner
helmets; two linkage assemblies mounted on the
cockpit left canopy frame; and an electronic
interface assembly mounted on the rear cockpit
bulkhead.

Aiming of TURRET is accomplished by super¬
imposing the reticle image on the target while
depressing the appropriate TRIGGER ACTION
switch. Error signals will cause the turret to move
until aligned with the viewer’s sight line. The
reticle image is projected by a reflex sight in
front of the operator’s right eye, and appears as a
yellow/white pattern focused at the target range.
Each linkage assembly (Pilot and Gunner) is
I stowed by sliding the respective linkage arm into
a spring loaded stow bracket at the forward end
of the linkage. In operation, the linkage arm is

connected to the pilot and gunner’s helmet by
means of a magnet at the rear of the helmet. This
attachment is for quick breakaway in the event
of an accident. Breakaway requires approximately
20 pounds of pull. Each linkage has a BIT (Built
in Test) magnet, to which the steel fastener at the
end of the linkage arms (both Pilot and Gunner)
are connected when performing BIT. The
respective linkage arm must be connected to the
helmet magnet to obtain a viewing reticle. The
turret weapons cannot be fired with the HS, if
this link-up has not been accomplished. The HS
reticle is supplied with light from one- of three
lamps. Three lamps are provided to ensure
reliability through redundancy. After linkage arm
link-up with the helmet, the HS reticle test switch
indicates failure of any of the three lamps. If one
lamp is inoperative, no reticle will appear when
the HSS RETICLE TEST switch is activated.

Each HS assembly has a cable terminating in an
eight-pin connector. This connector must be
attached to the mating jack located on the side
of the seat in the same clip as the communications
connector. After the HS assembly has been con¬
nected and the linkage attached, the sight eyepiece
must be positioned in front of the eye. To adjust
the eyepiece vertically, compress the ends of the
spring lock with the left hand and move the
eyepiece up or down with the right hand. To
adjust the sight laterally, grasp the sight housing
firmly and apply enough lateral force to overcome
the effect of the friction disc and cause the housing
to move. Tthe HS assembly can be retracted
(rotated out of the field of view) manually. To
retract the sight manually, push the small button
located to the right of the bulb cover. The sight
will rotate counterclockwise out of the field of
view. The signal for retracting the gunner sight
electronically occurs when the gunner moves the
ACQ/TRK/STOW switch from the ACQ position
to the TRK position or when the switch is in the
TRK position and the PHS ACQ switch is
depressed. The signal moves the sight out of the
way automatically before the gunner looks into
the TSU. The sight, once retracted, must be
manually returned to the field of view by rotating
the sight counterclockwise until it latches in front
of the eye. The turret is positioned in azimuth
and elevation by moving the HS with the linkage
attached, while the appropriate switch is depressed;
vertical movement of the head will produce eleva¬
tion movement of the turret and horizontal move¬
ment of the head will produce azimuth movement
of the turret. If the HSS assembly is on the
retracted position, positioning and firing circuits

8-17

Section VIII

NAVAIR 01-H1AAB-1

HELMET SIGHT SUBSYSTEM

PILOT/GUNNER LINKAGE ARM ATTACHMENT TO BIT MAGNET AND STOW BRACKET

PILOT/GUNNER LINKAGE ARM ATTACHMENT
TO HELMET SIGHT

1. Electronic Interface Assembly

2. Gunner Extension Cable

3. Pilot Linkage Cable

4. Pilot Linkage Arm

5. Pilot Linkage Rails

6. Pilot Helmet Sight

7. Pilot Eyepiece

8. Pilot Linkage Front Support

PILOT/GUNNER EYEPIECE
RETICLE PATTERN

9 Gunner Linkage Cable
10. Gunner Linkage Arm
11 Gunner Linkage Rails

12. Gunner Linkage Front Support

13. Gunner Helmet Sight

14. Gunner Eyepiece

15. BIT Magnet

1 6. Stow Bracket

210011-1

Figure 8-11. Helmet Sight Subsystem (HSS)

8-18

NAVAIR 01 -H1AAB-1

Section VIII

are not interrupted. Release of the action switch
will cause the turret to return to the stow position
regardless of HS position. If the action switch is
depressed and the sight is moved at a speed greater
than the maximum angular velocity of the turret,
the firing circuit is interrupted and the sight
reticle flashes until the gun is coincident within
5.5 degrees of the HS line of sight. The sight
reticle also flashes when turret is moved to
azimuth or elevation travel limits. HSS reticle
OFF-BRT knobs and TEST switches are located
on the pilot instrument panel and the gunner
armament control panel. An HSS BIT switch on
the gunner armament control panel tests the HSS,
PLT (Pilot), GNR (Gunner) and EIA (Electronic
Interface Assembly). Failure indicator lights plus
a GO (HSS operating properly) indicator light are
also located on the gunner armament control
panel.

TOW MISSILE SYSTEM (TMS) AH-1T/M65.

The tube launched, optically tracked, wire guided
(TOW) missile subsystem (TMS) is designed to
launch and guide the TOW antitank missile from
the AH-1T (TOW) Cobra. See figure 8-22 for (TOW)
missile system acronyms.

Functional Elements.

The TMS launching and guiding capabilities are
provided by five functional elements of hardware
packaged in eight Weapon Replaceable Assemblies
(WRAs), mounted within the aircraft. See figure
8-12 and FO-9.

The five functional elements are:

1. Stabilized sight.

2. Controls and displays.

3. Infrared.

4. Missile command.

5. Launcher.

The first element, the stabilized sight, provides the
capability for sighting and tracking a target using
commands generated by the gunner. This element
consists of the Telescopic Sight Unit (TSU) and
Stabilization Control Amplifier (SCA). An optical
telescope within the TSU allows the gunner to look
at a magnified image of the target for acquisition
and tracking. The TSU optics are stabilized to

effectively isolate the gunner’s field of vision from
the helicopter vibration and rotational motion.

The second element, controls and displays provide:
positioning commands to the TSU, system turn¬
on, missile selection, steering direction for the
pilot, system status to the gunner and pilot, self
test commands, and operational mode selections.
This element consists of the Sight Hand Control
(SHC), TOW Control Panel (TCP), Pilot Steering
Indicator (PSI), Status Annunciators and a Left
Hand Grip (LHG).

The third element, infrared (IR), provides the
capability of detecting the angular displacement of
the missile from the Optical Line of Sight (LOS),
by tracking an IR beacon which is located on the
aft end of TOW missile. The direction and ampli¬
tude of the angular displacement of the missile
from the TSU LOS is used to generate missile
position error signals. This element is located in
the TSU and consists of the IR tracker and error
detector.

The fourth element, missile command, processes
the missile position error signals that are generated
by the IR element into FM multiplexed signals.
These signals are transmitted over the missile
command wires as commands which are used to
direct the missile back to the TSU optical LOS.
This element consists of the Missile Control
Amplifier (MCA).

The fifth element, launcher, consists of the TOW
Missile Launcher (TML) attached to the outboard
wing stations (figure 8-13). The launchers are
designed so that either two or four missiles can
be loaded on each wing station. The interrelation¬
ship of the five elements is graphically illustrated
in figure FO-9.

Built In Test.

The TMS has an automatic BIT, consisting of ten
distinct tests designed to verify system operational
integrity and to indicate a failure of one of four
WRAs (TSU, SCA, MCA, and EPS). The EPS is the
only WRA in the sytem which maintains a
continuous operating integrity check and will give
a positive indication of malfunction any time the
TMS is receiving power. BIT is initiated whenever
the TMS is initially powered, and is immediately
sequenced, automatically to each test. When the
TMS is armed, BIT is interrupted and the system
is immediately ready for use. BIT cannot be
performed if ACQ/TRK/STOW switch is in any

8-19

Section VIII

NAVAIR 01-H1AAB-1

SIGHT

HAND CONTROL

PILOT

STEERING

INDICATOR

STABILIZATION

CONTROL

AMPLIFIER

MISSILE COMMAND
AMPLIFIER

ELECTRONIC

POWER

SUPPLY

Figure 8-12. Weapons Replaceable Assemblies

8-20

NAVAIR 01-H1AAB-1

Section VIII

FORWARD TUBE
ASSEMBLY

-FORWARD ATTACHING POINT
(lower rack use only)

— FORWARD ADJUSTABLE

BOMB LUG (upper rack use only)

SWAY BRACE PAD

AFT ADJUSTABLE

BOMB LUG (upper rack use only)

SWAY BRACE PADS

AFT ATTACHING POINT
(lower rack use only)

SWAY BRACE PAD

CENTER GATE
CAPTIVE LOCKING
PIN

FORWARD
ATTACHING POINT
(upper rack use on
when lower rack

installed) -

MISSILE ENGAGING
HANDLE

DEBRIS DIRECTOR
CAPTIVE LOCKING
PIN

HARNESS RECEPTACLE
(upper rack use only
when lower rack installed)

< HIDDEN >_DEBRIS

DIRECTOR
ASSEMBLY

HINGED CENTER GATE

AFT ATTACHING POINT
(upper rack use only
when lower rack
installed)

l— QUICK DISCONNECT
LANYARD

LAUNCHER

(LOOKING FORWARD)

209071-343

Figure 8-13. TOW Missile Launcher

8-21

Section VIII

NAVAIR 01 -HI AAB-1

position other than STOW. An indication appears
on the TOW control panel (TCP) that shows
whether BIT is In-Test (Test), pass (Power On)
or fail (OFF). If the system fails BIT, an
indicator will appear to isolate the failure to one
of the previously mentioned four WRAs. The BIT
equipment reset switch on the TOP resets the BIT
fail indicators and recycles BIT sequence. The
following units are not checked via BIT: TCP,
SHC, PSI and Launchers. When the manual BIT
button on the TCP is depressed, all annunciators
in the TSU and PSI will be displayed. When the
BIT button is released, the annunciators will
disappear from view and the BIT checks are
initiated.

Telescopic Sight Unit (TSU).

The TSU (figure 8-14) is one of the WRAs which
makes up the TMS. This WRA contains the optical
system necessary for firing the TOW missile.
Visually the TSU has an angular coverage of plus
and minus 110 degrees in azimuth and +30 to -60
degrees in elevation. The TSU is mounted on the
nose of the helicopter and extends into the front
cockpit. The SHC is mounted on the right side of
the TSU and the LHG is mounted on the left. The
IR tracker and error detector are located in front
of the TSU.

In operation, the IR tracker and error detector
receives IR energy from the missile during flight
and senses any missile displacement from the
optical LOS. This information is used to generate
the command signals which direct the missile back
to the LOS. During system self test, the IR tracker
is automatically boresighted to the optical LOS.
The stabilization system isolates the optical system
from helicopter motion.

| The optical fields of view offer a 28-degree field of
view in the 2x magnification, and a 4.6-degree
field of view in a 13x magnification. The gunner’s
left hand grip (LHG) is mounted on the TSU relay
tube in a location that optimizes the gunner’s
ability to grasp it for support and at the same time
operate necessary controls. The LHG controls are
used to select 2x or 13x magnification, provide a
trigger switch for firing the TOW missile, and
provide a weapon action switch for initiation of
the attack mode. Annunciators within the TSU
provide the gunner with system status information.
Mounted on the bottom of the TSU relay tube is
a focus knob which will compensate for any

astigmatism. The eyepiece is monocular and can
be rotated so that either eye can be used to view
through the optics.

Stabilization Electronic Control Amplifier (SCA).

The SCA is one of the WRAs making up the TMS.
It contains circuitry used to control power input to
the Electronic Power Supply (EPS). SCA motion
compensation circuitry has inputs from the SHC
track stick and TSU and provides azimuth/
elevation LOS drive output for the TSU. A servo
loop controls the TSU optics window and derota¬
tion error inputs. The servo loop generates window
and derotation servo drive signals to position the
TSU optics. The azimuth/elevation error signals
from the TSU and SCA drive signals are used to
generate pitch and yaw error signals and LOS
rates for the Missile Control Amplifier (MCA).
The yaw open loop command is generated from
gimbal resolvers and air speeds. The SCA contains
a constraints computer which determines whether
launch conditions are within pre-determined
boundaries. If the helicopter is within pre¬
determined boundaries, the constraints computer
provides a contraints valid signal to the MCA.
Stabilization circuits, error signal resolution,
motion compensation and open loop missile
steering are also provided by the SCA.

TOW Control Panel (TCP).

The TCP is the WRA which provides gunner
controls and indicators for TOW missile selection,
missile status, BIT, manual wire cut and camera
controls, if camera is installed. When the BIT is
depressed, ten sequential BITs are performed
automatically. The TCP receives a missile status
signal from the TML and displays missile status:
present (MSL), selected (SEL), missile not present
(BARBER POLE).

The TOW missiles can be selected manually (Mode
select: ARMED MANUAL or automatically (Mode
select: ARMED AUTO). If AUTO is selected, the
missile select switch will automatically sequence
to the first missile present by numbered priority
one through eight. If MAN is selected, each missile
must be selected manually. See figure 8-15.

Missile Command Electronic Control Amplifier
(MCA).

This WRA provides system timing, processes
missile guidance and command signals, and
provides automatic control of BIT.

8-22

Change 1

NAVAIR 01 -HI AAB-1

Section VIII

TELESCOPIC SIGHT UNIT
LOCATION: GUNNER STATION

13 X RETICLE FIELD OF VIEW IS 4.6°
FROM RETICLE CENTER TO EDGE OF

2 X RETICLE FIELD OF VIEW IS 28°

NOMENCLATURE

FUNCTION

Left Hand Grip Switches

Magnifies target two times.

MAG Switch

—

Magnifies target 13 times.

TRIGGER Switch

Press

—

Fires TOW in first or second detent.

Fires turret 16 ± 4 round burst in first detent.

Fires turret continuously in second detent.

ACTION Switch

Press

Activates TOW launchers.

—

Slaves turret to TSU or gunner HS.

LASER Switch

—

Inoperative.

TSU Reticle

GUNS Indicator

Flash

—

Indicates TCP MODE SELECT switch in
TSU/GUN and turret not aligned with

TSU.

Steady

—

Indicates TCP MODE SELECT switch in
TSU/GUN and turret aligned with TSU.

ATTK Indicator

—

Indicates TCP MODE SELECT switch in ARMED
and ACTION switch depressed.

RDY Indicator

—

Indicates pilot has achieved prelaunch constraints
for TOW firing.

Filter Select Lever

Move

—

Selects filters of different light intensities.

Focus Knob

Turn

—

Focus the target image.

Figure 8-14. Gunner Telescopic Sight Unit (TSU)

Change 1

8-23

Section VIII

NAVAIR 01 -HIAAB-1

NOMEN¬
CLATURE^ FUNCTION

MODE OFF
SELECT

Switch TSU/
GUN

Deactivates TSU and TMS
circuits.

Permits gunner to fire turret and
perform target acquisition, pilot to
fire smoke/wing stores (not TOW)
and perform target acquisition.
Activates built-in-test (BIT) when
SHC switch is in STOW, power
missile status indicators.

STBY -
TOW

ARMED-

MAN

ARMED-

AUTO

Permits gunner to control TMS/
perform target acquisition, pilot to
fire turret and perform target
acquisition. Activates built-in-test
(BIT) when SHC switch is in
STOW, power missile status
indicators.

Permits gunner to fire TOW
(manually selected) and perform
target acquisition, pilot to fire
turret, and perform target
acquisition.

Same as MAN except missile is
automatically selected.

NOMENCLATURE

TSU/SCA/

Black Flag

EPS/MCA

White Flag

Unit Fail indicators

Note

BIT Switch

Press

OFF

CAMERA

MAN

Switch

AUTO

EXPOSURE

BRT

Switch

HAZ

DUL

OFF/PWR ON/

OFF

ARMED/TEST

PWR ON

System Status

ARMED

Annunciator

TEST

TSU RTCL

OFF

Switch

Turn

WIRE CUT Switch

PRESS

MSL/Barberpole

MSL

Missile Status

Barberpole

Indicators

MISSILE

1/2/3/4/

SELECT Switch

5/6/7/8

Figure 8

FUNCTION

— Indicates unit operation during performance of built-in-test.

— Indicates unit failure after performance of built-in-test.

— Only EPS indicates failure at any time power is supplied
to the TMS.

— Performs manual built-in-test.

— Deactivates camera circuit.

— Permits continuous operation of camera.

— Permits operation of camera when LHG TRIGGER is pressed until missile
command wire is cut.

— Permits camera to adapt to bright conditions.

— Permits camera to adapt to haze conditions.

— Permits camera to adapt to dull conditions.

— Indicates MODE SELECT switch in OFF position.

— Indicates MODE SELECT switch in TSU/GUN or
STBY TOW position and BIT is complete.

— Indicates MODE SELECT switch in ARMED position.

— Indicates built-in-test is being performed.

— Deactivates the TSU reticle lamp.

— Varies intensity of TSU reticle lights.

— Permits gunner to manually cut missile command wire.

— Indicates missile is present in a specific location of launcher.

— Indicates missile is not present in a specific location of launcher.

— Indicates missile selected (manual or automatic) for firing.

Figure 8-15. Gunner TOW Control Panel (TCP)

21D071-39

8-24

NAVAIR 01 -HIAAB-1

Section VIII

MCA circuitry, provides automatic control of BIT
and BIT status signals to the TCP. TOW trigger
armed signal is used to start the MCA programmer
timer. The programmer provides a timing signal
in proper sequence for 23 seconds from initiation
of fire signal to automatic wire cut. The launch
constraints valid signal from the SCA or a
constraints override signal from the SHC is
required before the trigger in the LHG can initiate
the fire sequence. The open loop command from
the SCA is processed by the MCA and aids in
controlling the missile yaw during the interval
between launch and acquisition by the IR tracker.
The MCA utilizes pitch and yaw error signals, LOS
rates and G-bias signals to generate wire signals A
and B which are transmitted through the TML to
the missile. The MCA G-bias networks provide a
G-bias signal used to compensate for gravity on
the missile during missile flight. During interval
between initiation of fire sequence and launch
(1.5 seconds), self balance circuitry causes MCA and
missile frequencies to be aligned. A carrier network
generates signals to pitch and yaw channels pro¬
viding a duty cycle for the missile for its entire
flight.

Sight Hand Control (SHC) (Figure 8-16).

This WRA, mounted on the right side of the TSU
in the gunner’s station, contains the track stick
which is a force transducer device. It provides the
track commands which positions the TSU optics to
enable target locating and tracking. SHC controls
also select different modes of operation for the
TSU. The constraints override button allows the
gunner to initiate the TOW fire sequence if the
aircraft is not within constraints valid limitation.
(This function is normally only used by main¬
tenance personnel. Probability of missile capture
is greatly reduced utilizing constraints override).

Electronic Power Supply (EPS).

This WRA located in the tail boom, switches and
conditions aircraft power, provides power forms
not available from prime aircraft power sources,
and converts 28 vdc to provide regulated and
unregulated DC and AC voltage to operate the
TMS.

The EPS is energized by remote-on command from
the TCP. At system turn-on, BIT is performed
automatically. The EPS supplies its own BIT
signals and is the only WRA which maintains a
continuous BIT during operation. If during

operation, it fails, it will display the fail indication
on the TCP. A thermal switch, located in the
SCA, will activate if temperature limitations are
exceeded, causing the EPS to shut down, thereby
disabling the TMS.

TOW Missile Launcher (TML).

The TML is a WRA which provides the support
and electrical interface with the M-65 for up to 8
TOW missiles. The basic launcher (see figure 8-17)
holds two missiles. An expansion module identical
to TML can be mounted on the bottom of the
upper TML giving the aircraft the ability to carry
4 missiles on each outboard station. The lower
TML may be attached without repeating the
boresight procedure, provided the previously
boresighted upper TML is not removed. The upper
TML is supported by hooks beneath the aircraft
pylon and is retained/released by rotation of the
support hooks. The TML supports two remote
armament control boxes housing circuitry neces¬
sary to provide interface remote control to fire a
missile and isolation from the system prior to
firing. See figure 8-17 for missile location.

Pilot Steering Indicator (PSI).

The PSI is a WRA which assists the pilot in aligning
the aircraft within pre-launch constraints (±2.5°
AZ, ±6° EL, 5° angle of bank) in preparation for
missile firing and for maneuvering after firing
(±110 AZ, +60° to -30° EL, 30° angle of bank).
Two pointer type ascend/descend indicators will
indicate the direction to correct for excessive
pitch rates. See figure 8-18. The PSI has three
status annunciators. The ATTK annunciator tells
the pilot that all conditions are met for a TOW
launch except when the aircraft is not aligned in
pre-launch constraints. When the RDY annunciator
appears, all conditions are met for a TOW missile
launch and the fire sequence can be initiated by
the gunner with the LHG fire trigger. When the
gunner initiates the fire sequence, the FIRE
annunciator appears for the duration of missile
flight or until wire cut. The ATTK and RDY
annunciators will disappear from view at launch.
The only indication of aircraft attitude not
depicted by the PSI is angle of bank. However, if
the PSI indicates all aircraft alignment conditions
have been met and a RDY annunciator is not
visible, then the aircraft is probably out of
pre-launch roll constraints (±5°).

TOW Missile System Function.

(See inflight procedures all armament for
switchology.)

8-25

Section VIII

NAVAIR 01 -HI AAB-1

NOMENCLATURE FUNCTION

Track Control Stick

Move

— Positions TSU in aximuth and elevation.

ACQ/TRK/STOW

ACQ

— Slaves TSU to gunner HS for target acquisition.

Switch

TRK

— Permits track control stick to position TSU.

STOW

— Stows TSU dead-ahead.

PHS ACQ Switch

Press

— Slaves TSU to pilot HS for target acquisition.

EL BAL Screw

— Used during maintenance.

A2 BAL Screw

— Used during maintenance.

CONST OVRD Switch

Press

— Permits TOW firing when helicopter is not

aligned within the prelaunch constraint boundary.
(Normally used only for maintenance)

210071-40

Figure 8-16. Gunner Sight Hand Control (SHC)

8-26

NAVAIR 01 -HI AAB-1

Section VIII

LAU-68 Rocket Launcher

RIPPLE SINGLE

-o— O - o-O-©

Characteristics

Weight (Pounds)

(Empty) .

(Loaded).

Length (Inches) . .
Diameter (Inches) .
Suspension (Inches)

LAU-61, -69 Rocket Launchers

LAU-61, 69
Aft Fairing

Characteristics

Weight (Pounds)

(Empty) .

(Loaded)*
Length (Inches) .
Diameter (Inches)
Suspension

(Inches) .

*Mk 5 Warhead

LAU-61/A

. 132
.474
. 83.0 . .
. 15.7 . .

. 14.0 . .

LAU-69/A

98
. 440

. 83.0

. 15.7

. 14.0

NTS A 80

Figure 8-17. LAU Series Rocket Launcher (Typical)

8-27

Section VIII

NAVAIR 01 -HI AAB-1

LOCATION:

NOMENCLATURE

ATTK Annunciator

ON -

RDY Annunciator

ON -

FIRE Annunciator

ON -

Reference Ring
(Fixed)

—

Prelaunch

Constraint

Boundary (Fixed)

Postlaunch

Constraint

Boundary (Fixed)

Elevation/

Azimuth

Sightline Position
Bars (Moveable)

Ascend

Descend

ON -

Pointers

(Indicator)

OFF -

♦Azimuth Angle
Markers (Fixed)

—

♦Course Scale
Azimuth Pointer
(Moveable)

♦Fixed scale and gain.

Not affect

BASE OF PILOT FIXED SIGHT
FUNCTION

Indicates TCP MODE SELECT switch in ARMED position and TSU LHG ACTION
switch depressed.

Indicates pilot has achieved prelaunch constraints.

Represents boundary within which the pilot must keep the siahtline position bars Drior
to and during TOW launch. The boundary represents ± 2.5° azimuth and ± 6° elevation.

Represents boundary within which the pilot must keep the sightline position bars after
TOW launch and until cut or missile impact.

The boundary represents plus and minus 110° azimuth plus 60° to -30° elevation.

Indicates TSU elevation and azimuth angles with respect to helicopter reference axis
(reference ring) and constraint boundaries.

prelaunch constraints.

Indicates helicopter nose attitude and line-of-sight rate are compatible.

— Represents TSU 1110° azimuth limits.

Indicates TSU azimuth angle on the azimuth angle markers.

id by pre/post launch. 210071-3 7

Figure 8-18. Pilot Steering Indicator (PSI)

8-28

NAVAIR 01 -HIAAB-1

Section VIII

Target Acquisition.

The gunner has several methods of acquiring a
target through the TSU. By placing the ACQ/TRK/
STOW switch in the TRK position. The TSU may
be directed towards the target utilizing the SHC
stick. This method will take more time than other
methods and will require the gunner to search with
the TSU. As the field of view is restricted to 28
degrees in LO mag and 4.6 degrees in HI mag,
some difficulty may be encountered. The quickest
method is to utilize either the PHS or GHS to
direct the TSU optics. If the pilot places the PHS
reticle on a target, the gunner can direct the TSU
to that target by placing the ACQ/TRK/STOW
switch on the SHC to TRK and depressing the PHS
ACQ button, also located on the SHC. When the
PHS ACQ button is depressed, the GHS reticle
automatically retracts enabling the gunner to view
through the TSU. The TSU will continue to align
with the pilot’s LOS until the gunner releases the
PHS ACQ button.

If the gunner desires to direct the TSU to a target
using the GHS, he proceeds as follows:

1. Superimpose the GHS reticle on target.

2. Move the ACQ/TRK/STOW switch to ACQ.

The ACQ/TRK/STOW switch is spring loaded from
ACQ to TRK so it will be necessary to hold it in
the ACQ position. As long as the switch is held in
the ACQ position, the TSU will continue to align
itself with the GHS LOS. When the switch is
released, it will spring back to the TRK position
and the GHS sight will automatically retract. The
gunner then views through the TSU to re-acquire
the target. The acquisition functions will operate
for any mode select position on the TCP except
OFF.

TOW Missile Firing.

Once the gunner has acquired a target in the LO
mag, he switches to the HI mag position on the
LHG. The small circle in the LO mag reticle of
the TSU represents the limits of the HI mag field
of view. If the target appears in the small circle
of the LO mag reticle, the target will appear
within the HI mag field of view. When in HI mag,
the gunner should keep the weapons action switch
depressed to get motion compensation and to
complete the attack logic necessary to launch the
TOW missile.

Weapons action switch - DEPRESSED (motion
compensation)

1. HI mag - SELECTED

2. ACQ/TRK/STOW switch - TRK

3. Missile - PRESENT and SELECTED

4. TCP Mode Select - ARMED MAN or AUTO.

With these conditions met, the ATTK annunciator
will appear in the TSU field of view and the PSI.

The PSI gives the pilot steering information to
align the aircraft within pre-determined pre-launch
constraints. As the aircraft comes into pre-launch
constraints, the RDY annunciator appears on the
PSI and within the TSU field of view. The pilot
should strive to give the gunner as stable a platform
as possible for the actual firing.

With the ATTK and RDY annunciators present,
the firing sequence can be initiated by the gunner
utilizing the TRIGGER on the LHG. By pulling the
LHG TRIGGER and initiating the fire sequence, a
fire annunciator appears on the PSI and the ATTK
and RDY annunciators will disappear from view at
launch. No annunciators will be evident in the
TSU. After initiation of the fire sequence, there
will be a 1.5 second delay before the missile
launch. The 1.5 second delay is necessary for the
following:

1. Missile battery charge-up.

2. Missile gyro spin-up.

3. Missile guidance set self-balance.

1.5 seconds after initiation of the fire sequence,
the launch motor ignites. The launch motor
accelerates the missile to 225 FPS and the missile
holdback pin is sheared allowing the missile to
exit the launch container. The launch motor
bums out before the missile exits the launch tube
and the missile coasts approximately 7-12 meters
before the flight motor ignites. At this point, the
wing and flight surfaces have snapped out into
position and the flight motor ignites accelerating
the missile to just under Mach 1. When the flight
motor ignites, the acceleration of the missile causes
a G-sensing device to complete the missile arming.
At this point, nose crush is all that is necessary
to detonate the warhead. The flight motor quickly
bums out and the missile coasts for the duration of

Change 1

8-29

Section VIII

NAVAIR 01 -HI AAB-1

I the flight. Ignition of the flight motor will cause
| target obscuration, due to smoke and gases, for a
short period of time. As target obscuration occurs,

I the gunner should release his control inputs with
the SHC and allow motion compensation to keep
the STU crosshairs on the target. As obscuration
decreases, if the crosshairs have drifted off the
I target, the gunner should make a smooth positive
correction back to the target avoiding jerky SHC
movement.

When the missile is fired, the gains on the PSI
change to postlaunch constraints and the maneuver
limits are now represented by the large boundaries
on the PSI. The pilot should strive to minimize
| erratic aircraft movement to minimize gunner
tracking error.

The gunner continues to track the target until
missile impact or wire cut. Wire cut will be auto¬
matically initiated by missile impact or by timer
23 seconds after TRIGGER pull or if IR tracker
loses the missile or source for more than .5
seconds. Manual wire cut can be initiated by either
the gunner or pilot, at any time, utilizing the
respective wire cut buttons (see FO-6 and FO-7).

Because of the nature of the missile
flight controls, when wire cut occurs the
missile flight will be extremely erratic.

TSU GUNS.

With TSU guns selected on the TCP mode
SELECT, the gunner has the ability to fire the
20-mm turret in the flex mode utilizing the
stabilized optics of the TSU. The ACQ/TRK/
STOW switch must be in the TRK position and the
LHG controls are utilized for weapons action and
fire. The LO mag position should be utilized for
the first firing burst. If the impacts are within the
small circle of LO mag reticle, then the impacts
will appear in the HI mag field of view. If the
impacts are out of the small LO mag circle but,
in a vertical plane with it, then an adjustment on
the RANGE knob on the gunner armament control
panel may bring the impacts into the small circle.
If the impacts are horizontally out of the small
LO mag circle, range adjustments will not bring
the impacts into the HI mag field of view and HI
mag should not be selected. Depression of the
weapons action bar aligns the gun barrels with the

TSU LOS and in HI mag also gives motion compen¬
sation. The TSU/GUNS TRACK RATE switch on
the gunner’s armament control panel, gives the
gunner the ability to select a HI or LO track rate
for the TSU while in TSU/GUN.

If the gunner fails to place the ACQ/
TRK/STOW switch in STOW when
coming out of the TSU and attempts
to fire the GHS, the turret will fire in the
direction of the TSU LOS.

Gunner Accuracy Control Panel (GACP).

The Gunner Accuracy Control Panel is a training
device which allows the TMS to be operated under
simulated conditions and is depicted in figure 8-19.
A ground IR source is necessary (M-70) to utilize
the GACP. The TMS must be armed for the GACP
to function. After the system is armed and the
GACP turned on, it will self test. The digital
indicators will display 88, 99, and then return to
55±2 after 14 seconds. No calibration is necessary.
The TSU must be receiving an IR signal from the
M-70 before a firing sequence can be initiated.
The TMS can be operated as though a missile were
to be launched. The GACP provides a missile
present indication and missile 5 must be selected.
The missile present signal will disappear after the
fire sequence has been initiated but will return
after the scoring run has been completed. The
azimuth and elevation meters provide a visual
indication of tracking errors so the pilot can
critique the gunner during the 12-second scoring
run. It is not necessary to reset the GACP after
the scoring run is completed. The PSI will display
the FIRE indicator during the 12-second scoring
sequence. The GACP panel will present a score
for both azimuth and elevation after the scoring
sequence is complete. The GACP is normally
installed only for training purposes and is mounted
on a bracket to the right side of the pilot’s rocket
sight.

A TOW missile shall not be fired with
GACP installed due to erratic missile
response.

8-30

NAVAIR 01 -HI AAB-1

Section VIII

LOCATION: TOP OF PILOT INSTRUMENT PANEL IF INSTALLED

NOMENCLATURE
ON/OFF Switch

WIRE CUT Switch
AZIMUTH Indicator
AZIMUTH SCORE Indicator

ELEVATION Indicator
ELEVATION SCORE Indicator

BRIGHTNESS Knob

FUNCTION

ON — Activates gunner accuracy control circuits.

— Performs built-in-test of circuits. Circuits pass test if
AZIMUTH/ELEVATION SCORE indicators display 55
+ 2. (TMS has to be ARMED.)

OFF — Deactivates circuits.

Press— Resets GACP and deactivates camera.

— Displays TSU azimuth line-of-sight deviation.

— Displays gunner final azimuth score.

_ Displays TSU elevation line-of-sight deviation.

— Displays gunner final elevation score.

Turn — Varies intensity of AZIMUTH/ELEVATION SCORE
indicator lights.

210071-41

Figure 8-19. Pilot Gunner Accuracy Control Panel (GACP)

8-31

Section VIII

NAVAIR 01 -HI AAB-1

WING STORES JETTISON.

The pilot has an EMERGENCY JETTISON
SELECT panel on the pilot instrument panel and
a guarded JETTISON switch on the collective
switch box (FO-6). The gunner has a WING
STORES JETTISON switch on his armament
control panel and a guarded JETTISON switch
on the gunner miscellaneous panel (FO-7).
Activation of a jettsion switch (pilot or gunner)
fires the cartridges and separates the stores. Pilot
jettison switch must be held depressed for at least
one second to assure cartridge firing. Circuits
are protected by the WING STORES JTSN PLT
and WING STORES JTSN GNR circuit breakers on
the pilot AC/Armament circuit breaker panel.

WING STORES ARMAMENT SYSTEM.

Four attachment points are provided, two under
each wing. The pylon assemblies include external
store racks, sway braces and standard electrical
connections for external stores. The entire
assembly is enclosed in a fairing that matches the
lower contour of the wing.

The ejector rack of each pylon is equipped with an
electrically operated ballistic jettison device. The
jettison system consists of a breech block that
utilizes cartridges with independent firing circuits.

ROCKETS.

The 2.75-inch folding fin aerial rocket (FFAR)
subsystem is a light anti-personnel /assault weapon.
A launcher (figure 8-17) can be mounted on each
inboard and outboard ejector rack.

WING GUN POD (GPU-2A).

The self contained pod (figure 8-20) houses a
20-mm machine gun, electrical system, battery
recharging system and has a capacity of 300
rounds of ammunition. The gun is capable of
firing 750 rounds per minute.

SMOKE GRENADE DISPENSER (M-118).

A dispenser (figure 8-21) may be attached to each
outboard ejector rack. Each dispenser contains two
independently operated racks of six white or
colored smoke grenades, 12 per dispenser. One to
four grenades may be dropped at one time by the
two dispensers.

NOTE

Interference by nuts and bolt fasteners
on the body of the M-118 will prevent
the sway brace bolt pads from complete
seating on the dispenser.

PREFLIGHT PROCEDURES.

Before Exterior Check -All Armament - Preflight

• Personnel should remain clear of gun
and turret travel area when helicopter
electrical circuits are energized.

• Personnel should remain clear of
hazardous area of loaded weapons.

• Helicopters with loaded weapons should
be pointed toward clear area.

1. MASTER ARM — OFF

2. PILOT OVERRIDE — OFF

3. ALE-39 ARM — SAFE

4. ALE-39 POWER — OFF.

Exterior Check - Preflight.

M 197 GUN SYSTEM.

1. Gun barrels - CHECK FREE ROTATION

2. Ammunition — VISIBLE IN FEEDER

3. Safing solenoid — DISCONNECTED

4. Elevation brake - ON (down position)

5. Turret fairing and access doors —INSTALLED
AND SECURE.

WING STORES.

1. Wing ejector racks safety lever — LOCKED

2. Detent safety pins - INSTALLED

3. Check pods for security — SHAKE EACH
END

8-32

Change 1

NAVAIR 01 -HI AAB-1

Section VIII

Figure 8-20. Wing Gun Pod (GPU-2A)

8-33

Section VIII

NAVAIR 01-H1AAB-1

Figure 8-21. Smoke Grenade Dispenser

TOW

—

Tube launched, optically tracked,
wire guided

TMS

—

TOW Missile System

TSU

—

Telescopic Sight Unit

TCP

—

TOW Control Panel

MCA

—

Missile Control Amplifier

SCA

—

Stabilization Control Amplifier

SECU

—

SERVO Electronic Control Unit

EPS

—

Electronic Power Supply

SHC

—

Sight Hand Control

PSI

—

Pilot Steering Indicator

BIT

—

Built In TEST

WRA

Weapons Replaceable Assembly

TML

—

TOW Missile Launcher

EIA

—

Electronic Interface Assembly

LHG

—

Left Hand Grip

HSS

—

Helmet Sight Subsystem

PHS

—

Pilot Helmet Sight

GHS

—

Gunner Helmet Sight

LOS

—

Line-Of-Sight

—

Infrared.

Figure 8-22. TOW Missile System Acronyms

8-34

NAVAIR 01 -HI AAB-1

Section VIII

4. Jettison cable - CONNECTED AND CART¬
RIDGE INSERTED

5. Rocket cable - DISCONNECTED FROM
POD

6. Jettison feet - FLUSH WITH PODS.

7. ALE-39 DISPENSER POD SAFETY
switches — SAFE.

MK 81 /MK 82 BOMBS.

1. Fuze /fuze extension HANDTIGHT, Arming

delay — SET

2. Fuze arming assembly; arming vane and 2
Fahnestock clips — INSTALLED

3. Arming wires — NOT PRELOADED

4. Arming wire (M904E2/E3/E4) PROPERLY
ROUTED; 3 Fahnestock clips — INSTALLED

5. Single (Mk 9) arming wire (M1A1 Fuze
Extension) — PROPERLY ROUTED,. 3
Fahnestock clips INSTALLED; Arming
wire taped to fuze extension

6. Arming wires to rack arming solenoids —
ATTACHED; Fahnestock clip — INSTALLED

7. Fuze and arming device/wires — REMOVED

8. Fin release wire — INSTALLED (if appli¬
cable); Safety pin — REMOVED

9. Overall condition — CHECK.

CBU-55 FUEL AIR EXPLOSIVE.

1. FMU-83/B fuze — INSPECT

a. Fuze cover — REMOVED

b. Fuze delay — SET

c. Fuze safety pin — REMOVED

2. Tail fin thruster safety pin warning
streamer — REMOVED

3. Arming wire extractor — CONNECTED

4. Fins - SPREAD/LOCKED

5. Leak detector — SAFE

6. Overall condition — CHECK.

MK 77 MOD 2/4 FIRE BOMBS.

1. Fire bomb — NO LEAK OR DAMAGE

2. Fuze (AN-M173A1/M918) - WRENCH
TIGHT

3. Initiator (Mk 13) — INSPECT

a. Retaining rings — TIGHT

b. Tear-out section — NOT DAMAGED

c. Functioning delay — SET

4. Arming wires/lanyards — NOT PRELOADED

5. Arming wire (AN-M173A1/M918)

ATTACHED

6. Arming wires/lanyards — ATTACHED

7. Fuze and arming wire device safety pins/
wires — REMOVED (If applicable).

GPU-2/A GUN POD.

1. Helicopter adapter cables — DISCONNECT

2. Fire volts access doors — OPEN

3. Battery cable - DISCONNECTED

4. Drum — LOADED

5. Overall condition — CHECK.

FLARE DISPENSER.

SUU-44/A.

1. Detent safety pins — INSTALLED

2. Breech caps — HAND TIGHTENED

3. Spider cables — CONNECTED

4. Shear pins — INSTALLED FROM TOP,
ENDS BENT

5. Dispenser — OVERALL CONDITION.

Change 1

8-35

Section VIII

NAVAIR 01-H1AAB-1

MK 45 FLARE.

1. Ejection dial — SAFE

2. Split ring — RED DISC INSTALLED

3. Swivel snap hook — NOT ATTACHED.
SMOKE GRENADE DISPENSER.

Ml 18.

1. Electrical harness — CONNECTED

2. Grenade safety pin — IN PLACE

3. Grenade fuze — SECURE

4. Note color and position of installed grenades
(needed for interior check).

ROCKET LAUNCHERS.

LAU 61/68/69.

1. RADHAZ shield - SECURE (if required)

2. RIPPLE/SINGLE MODE Selector switch —
SINGLE .

3. Select dial — ARM

4. Launcher — OVERALL CONDITION.

TOW.

1. Launcher Mounting — Upper launcher aft
and forward adjustable bomb lugs secure to
helicopter ejector racks and racks swaybrace
bolts firmly against launcher swaybrace pads.
Lower launcher aft and forward attaching
points secure to upper launcher aft and
forward attaching points.

2. Electrical connector — Upper launcher
harness connected to helicopter receptacle
and jettison quick disconnect lanyard
attached to harness and launcher.

Lower launcher harness connected to upper
launcher harness receptacle.

3. Missile Installation — Missile container front
ring seated in forward tube mating ring,
hinged center gate and debris director secure
with captive locking pins. Note number of

Section VIII

and position of installed missiles (needed for
interior check).

ARMING/DEARMING PROCEDURES - IN
ARMING AREA.

1. Appropriate arming heading — ASSUME

2. Throttles - OPEN

3. Armament circuit breakers — DE-ENER-
GIZED

4. Smoke grenade switches — OFF

5. Emergency jettison select switches — OFF

6. Canopy jettison pins — IN

7. MASTER ARM — OFF

8. WEAPON CONT — FIXED

9. NARCADS store control panel —

a. Bomb arm — SAFE

b. QTY - 0/0

10. ALE-39

a. ARM switch — SAFE

b. PWR switch — OFF I

11. Gunner armament control panel —

a. PILOT OVERRIDE - OFF

b. TURRET DEPR LIMIT - LIMIT

c. WING STORES SELECT — OFF

12. TCP MODE SELECT — OFF

13. ACQ /TRK/STOW switch — STOW

14. RAD ALT —OFF

15. TACAN — RECEIVE

16. IFF — STBY

17. HANDS - IN VIEW OF ORDNANCE
PERSONNEL.

8-36

Change 1

NAVAIR 01-H1AAB-1

No radio transmissions shall be made
within 50 feet of arming or dearming
aircraft.

AFTER ARMING

1. RAD ALT/TACAN - ON

2. IFF - AS REQUIRED

8 3. WEAPON CONTROL, HSS, JETTISON circuit

breakers - ENERGIZED

4. Rounds remaining — SET

5. PHS and GHS rail arm assembly - ATTACH
TO BIT MAGNET

6. MASTER ARM - STBY

7. Weapons control — GUNNER

8. HSS BIT — BIT/RELEASE (BIT will complete by
2.5 seconds)

9. PHS and GHS rail arm assembly — ATTACH
TO HELMETS

10. PHS and GHS sight assembly — ADJUST

11. Pilot and gunner HSS test — TEST
| 12. TSU guns - TRACK RATE switch — HIGH

13.TCP MODE SELECT — STBY TOW

When TCP status indicator indicates test:

14. TCP BIT button — DEPRESS (verify all TSU/

PSI indicators appear)

15. TCP BIT button - RELEASE

After BIT is complete, which can take up to 2
minutes:

16. TCP status indicator — INDICATES POWER
' ON

17. TCP BIT status indicators — INDICATE ON

18. Jettison select — AS REQUIRED

19. Take-off checklist — COMPLETE

INFLIGHT PROCEDURES - ALL ARMAMENT.

The following armament inflight procedures para¬
graphs are based on only one weapon installed, all
armament circuit breakers in pilot RECOIL COMP,
switch on. Refer to FO-8 for firing modes when
two or more weapons are installed.

Do not engage cyclic or LHG switches
during any switching action on arma¬
ment control panels.

, WWMt WWM M WMW * ,

CAUTION

• If weapon firing stoppage occurs, imme¬
diately release firing switch or extensive
damage to equipment may occur. Do not
attempt to fire weapon until stoppage
corrective action has been taken.

# In the event of runaway gun, place
MASTER ARM/PILOT OVERRIDE
switch OFF.

Turret Operation.

GUNNER OPERATION — TURRET.

1. Pilot MASTER ARM — ARM

2. Pilot WEAPON CONT - GUNNER

3. Pilot RECOIL COMP - ON

4. Gunner PILOT OVERRIDE - OFF

5. Gunner RANGE — AS DESIRED

6. Gunner AIRSPEED COMP - COMP

7. Gunner TURRET DEPR LIMIT - OFF

8. Gunner to use HS

a. TCP MODE SELECT — OFF

b. TCP MODE SELECT - TSU/GUN

Change 1

8-37

NAVAIR 01-H1AAB-1

Section VIII

c. ACQ/TRK/STOW - STOW
or

d. PILOT OVERRIDE - OVERRIDE

e. Cyclic turret switches — UTILIZE VICE
LHG TURRET SWITCHES

9. Gunner to use TSU/GUN

a. TCP Mode Select — TSU/GUN

b. ACQ/TRK/STOW--TRK

c. VIEW through TSU to fire

10. Gunner LHG ACTION — DEPRESSED

11. Gunner HS/TSU recticle — ON TARGET

12. Gunner LHG TRIGGER — DEPRESSED.

First detent 16 ±4 round burst, second

detent continuous.

NOTE

The pilot can interrupt firing by deener¬
gizing the appropriate circuit breaker,
or depressing the cyclic TRIGGER
ACTION Switch.

| PILOT OPERATION - - TURRET.

1. Pilot MASTER ARM — ARM

2. Pilot RECOIL COMP — ON

3. Gunner PILOT OVERRIDE - OFF

4. Pilot RANGE — AS DESIRED

5. Gunner AIRSPEED COMP — COMP

6. Gunner TURRET DEPR LIMIT — OFF.

| When using HS:

1. Pilot WEAPON CONT — PILOT

2. Pilot cyclic TRIGGER ACTION —
DEPRESSED (If not depressed gun will fire -
from lower stow.)

4. Pilot cyclic TRIGGER TURRET FIRE —
DEPRESSED
First detent 16±4 round burst, second detent
continuous.

When using fixed sight:

1. Pilot WEAPON control — FIXED

2. Pilot sight - SET

3. Pilot sight reticle — ON TARGET

4. Pilot cyclic TRIGGER ACTION —
DEPRESSED.

5. Pilot cyclic TRIGGER TURRET FIRE —
DEPRESSED.

First detent 16±4 round burst, second detent
continuous.

NOTE

With weapon control in gunner and TCP
mode select in a TOW mode, the pilot
will have control of the turret and can
fire from lower stow or utilize his HS.

When the gunner comes out of the TSU,
the LHG mag switch should be placed in
LO mag and ACQ/TRK/STOW to STOW
to prevent firing of system in undesired
mode.

TOW Operation.

1. MASTER ARM — STBY

2. Pilot WEAPON CON — GUNNER

3. TCP MODE SELECT - ARMED MAN
OR AUTO

4. MISSILE SELECT — SET AND INDICATES
SEL

5. ACQ/TRK/STOW — TRK

6. PHS ACQ — DEPRESS FOR PHS ACQ THEN
RELEASE WHEN TARGET ACQUIRED

7. ACQ/TRK/STOW - ACQ FOR GHS ACQ
THEN RELEASE

8-38

NAVAIR 01-H1AAB-1

Section VIII

8. LHG ACTION - DEPRESS

9. MAG switch - HI AFTER TARGET IS IN HI
MAG FIELD OF VIEW

10. AIRCRAFT - MANEUVER INTO PRE¬
LAUNCH CONSTRAINTS AS INDICATED
BY PSI

11. MASTER ARM - ARM

12. LHG TRIGGER - DEPRESS

13. LHG TRIGGER - RELEASE

14. MASTER ARM - STBY.

LHG TRIGGER shall be released after
missile launch to prevent inadvertent
firing of next missile.

Jettisoning of TOW launchers for a
misfire condition is extremely dangerous;
do not jettison launcher unless fire is
encountered.

NOTE

• Smoke may emerge from launcher
after TRIGGER is depressed and before
missile exits launcher. Smoke is caused
by missile gyro and battery squibs firing
and should not be regarded as a misfire.

• If TOW missile fails to exit from the
launcher within 1.5 seconds and the PSI
RDY annunciator disappears, a misfire
has occurred.

• Gunner may attempt second firing by
releasing and depressing LHG TRIGGER;
if TOW missile again fails to fire, set
TCP, MISSILE SELECT switch to
select another TOW missile.

• Gunner cannot fire if helicopter is not
within prelaunch constraints boundary.
Gunner can override prelaunch constraint
boundary limitation by pressing CONST
OVRD switch on the SHC. If this mode
of operation is employed, degraded
system performance can be expected.

15. Helicopter — MANEUVER. Keep pilot PSI
sightline position bars within postlaunch
constraint boundary until wire cut or missile
impact.

NOTE

• Loss of missile guidance could result if
postlaunch constraints are exceeded.

• Missile wires are cut automatically and
PSI FIRE annunciator disappears. If
wires are not automatically cut, the pilot
can manually cut the wires using the
PILOT WIRE CUT switch or gunner can
manually cut wires using TCP WIRE
CUT switch.

Additional missile firing — Next missile is selected
automatically when gunner TCP MODE SELECT
IN ARMED AUTO; manually in ARMED MAN.

Rocket Operation.

1. Pilot MASTER ARM - ARM

2. Gunner PILOT OVERRIDE - OFF

3. Pilot STATION SELECT - SELECT

4. Pilot RATE — AS DESIRED

5. Pilot QTY - AS DESIRED

6. Pilot MODE - AS DESIRED

7. Pilot cyclic WING ARM FIRE
DEPRESSED.

Wing Gun Pod Operation.

1. Pilot MASTER ARM - ARM

2. Gunner PILOT OVERRIDE - OFF

3. Pilot STATION SELECT - SELECT

4. Pilot sight - ON TARGET

5. Pilot cyclic WING ARM FIRE —
DEPRESSED.

Smoke Grenade Dispenser Operation.

1. Pilot MASTER ARM - STBY or ARM

2. Pilot LH and RH ARM — AS DESIRED

8-39

Section VIII

NAVAIR 01-H1AAB-1

3. Pilot SMOKE RELEASE — DEPRESS.
Bomb Operation.

1. Pilot MASTER ARM — ARM

2. Gunner PILOT OVERRIDE — OFF

3. Pilot STATION SELECT - SELECT

4. Pilot BOMB ARM — AS DESIRED

5. Pilot cyclic WING ARM FIRE —
DEPRESSED.

Flare Operation.

1. Pilot MASTER ARM - ARM

2. Gunner PILOT OVERRIDE - OFF

3. Pilot STATION SELECT - SELECT

4. Pilot cyclic WING ARM FIRE —
DEPRESSED.

POST FIRING/BEFORE LANDING CHECK -
ALL ARMAMENT.

1. TCP MODE SELECT - OFF

2. PILOT OVERRIDE - OFF

3. TURRET DEPR LIMIT - LIMIT

4. MASTER ARM - OFF

5. WEAPON CONTROL - FIXED

a. ARM switch — SAFE

b. PWR switch — ARM

6. Armed/CP lights - EXTINGUISHED

7. Turret stow lights — ILLUMINATED

8. Armament circuit breakers —
DE-ENERGIZED.

After Dearm.

1. RADALT/TACAN/IFF — AS DESIRED

2. Take-off checklist — COMPLETE

AN/ALE-39 COUNTERMEASURES DISPENSING
SYSTEM.

xiie countermeasures Dispensing |

System (figure 8-23) permits the pilot or copilot
to selectively eject flares, chaff, or active radio
devices (jammers) from dispensing pods on the
stub wings. These items are designed to defeat
enemy surveillance radar, missile guidance radar,
and passive homing missiles. The AN/ALE-39 has
the capability of dispensing up to sixty chaff, flare,
and jammer payloads loaded in any combination in
multiples of ten. All three types of payloads can be
dispensed in both manual (single) and automatic
(programmed) modes independently or
simultaneously. The dispensing function can be
initiated by the pilot, copilot, or a radar warning
receiver system. The AN/ALE-39 system consists
of two dispenser housings, two dispenser
assemblies, two pilot and copilot actuator switches,
two sequencer switch assemblies, one programmei
assembly, one ALE-39 control panel, and one
ale 39 arm voltage control.

Countermeasures Dispensing System Operating
Procedures.

1. DISP and CONT cricuit breaker — IN

2. ALE-39 PWR switch — ON

3. ALE-39 ARM switch — ARM

4. MODE SEL — AS REQUIRED

5. Pilot/copilot DISPENSER switch - AS
REQUIRED.

AN/APR-39 RADAR WARNING SYSTEM.

The AN/APR-39(V) 1 is a passive omnidirectional
radar warning system receiving and displaying
information to the pilot concerning the radar |
environment surrounding the aircraft. The equip¬
ment responds to radar signals associated with
hostile fire control radar in E, F, G, H, I and J
frequency bands (wide-band) and provides visual
and aural indications of the presence and direction
of emitters. Radar signals which are not hostile are
generally excluded.

8-40

Change 1

NAVAIR 01-H1AAB-1

Section VII

Missile guidance radar signals in C and D bands are
also received by this system. When a low-band
signal is correlated with a tracking radar signal, the
equipment identifies the combination as an acti¬
vated SAM radar complex. This system consists of
four spiral antennas, one blade antenna, a
comparator, an APR-39 control panel, two
receivers and an APR-39 radar signal indicator.

The control panel (figure 8-24) is located on the
right side of the pilot’s glare shield. System control
and test functions are provided by this unit.

Radar Warning System Operating Procedures.

1. Radar WRN circuit breaker — IN

2. DSCRM switch - OFF

3. PWR-ON/OFF switch - ON

Allow a minimum of 30 seconds for equip¬
ment to become fully operational.

4. Audio — ON

5. IFF-Intercom switch — ON

6. Self-test switch — DEPRESS

a. Adjust BRIL and filter

b. Adjust AUDIO to desired level

7. DSCRM switch - AS REQUIRED.

To prevent damage to the receiver
detector crystals, ensure that the AN/
APR-39 antenna are at least 60 yards

from active ground based radar antenna,
or 6 yards from active airborne radar
antenna. Allow an extra margin for new,
unusual, or high powered antennas.

AN/ALQ-144 COUNTERMEASURES SYSTEM.

The AN/ALQ-144 is an active countermeasures
system which provides mechanical modulation of
radiation from an electrically heated source designed
to defeat the homing of approaching hostile heat
seeking missiles. The system consists of a trans¬
mitter, an operator control unit, and a bus transfer
relay assembly. The operator control unit (figure
FO-6) is positioned at the bottom of the armament
control panel between the pilot’s legs. Control,
operating test, and display functions are provided
by this unit. Control is provided by an ON/OFF
switch (figure 8-25) which activates the system
by applying 28 vdc power.

IR Jammer Operating Procedure.

1. IR JAMMER (XMTR, CONT, BASE) circuit
breakers — IN

2. ON/OFF switch - ON

If an INOP condition is indicated by
illumination of the IRCM light located
adjacent to the pilot’s sight, de-energize the
transmitter by setting the ON/OFF switch to
OFF.

Change 1

8-41

Section VIII

NAVAIR 01-H1AAB-1

NOMENCLATURE

FUNCTION

1. Power switch

2. ALE-39 Arm switch

3. Arm light

4. Counters

OFF — Power off to ALE-39

ON — Activates ALE-39

Salvo Flare — Fires all flares in dispensers

OFF - Disables ALE-39

ON — Master arm switch for ALE-39

Extinguished when Switch #2 is OFF or when Switch #2
is on dispensers installed. Armed light is on when switch ON
and ARM pin removed.

Indicates sumber of payloads, by type, remaining in dispensers.

C - chaff
F — flare
J — rf jammer

5. Payload reset

Sets quantity of each type of payload loaded.

N2/83

Figure 8-23. Countermeasures Dispensing System (Sheet 1 of 2)

8-42 Change 1

NAVAIR 01-H1AAB-1

Section VIII

NOMENCLATURE FUNCTION

6. Mode Select switches (one for
each countermeasure)

0 — Disables that countermeasure

S — Single, one countermeasure per actuation of pilot jettison
switch

P — Program, initiates program sequence as per programmer

R — RWR, series of payloads will be dispensed under control
of the radar warning receiver

M - Multiple, burst of 2, 3 or 4 flares in parallel, depending on
the number of dispenser sections containing flares

G - Group, multiple bursts of flares dispensed as per pro¬
grammer

7. Dispenser switches pilot/copilot

Push to dispense

Push to initiate dispense sequence

8. B QTY switch

CHAFF Section

1, 2, 3, 4, C, or R selects number of chaff bursts in one salvo
(C is continuous and R is random).

9. B IN TV switch

1, 2, 5, 7, 10 or R selects time interval between chaff bursts
of each salvo in seconds (R is random).

10. S QTY switch

1, 2, 4, 8, 10, or 15 selects number of chaff salvos required to
end programmed sequence.

11. S INTV switch

2, 4, 6, 8, or 10 selects time interval between chaff salvos in
seconds

FLARE Section

12. QTY switch

2, 3, 4, 6, 8, or 10 selects number of flare bursts required to
end flare programmed sequence.

13. INTV switch

2, 4, 6, 8, or 10 selects time interval, in seconds, between
bursts in programmed sequence.

LOAD Section

14. L10 switch

C, F, or J indicates type of payload in L10 dispenser.

15. L20 switch

C, F, or J indicates type of payload in L20 dispenser.

16. R20 switch

C, F, or J indicates type of payload in R20 dispenser.

17. RIO switch

C, F, or J indicates type of payload in R10 dispenser.

18. RESET switch

When pressed (3 seconds minimum) clears all registers and
counters in programmer and resets sequencer switches.

JAMMER Section

19. INTV switches

Selects in seconds the time interval between bursts of pro¬
grammed sequence (continuous from 000 thru 299).

20. QTY switch

1, 2, 3, or 4 selects number of jammer bursts required to end
programmed sequence.

♦Refer to NAVAIR 16-30A39-1, Intermediate Maintenance Manual, for a discussion of programmer operation.

N2/83

Figure 8-23. Countermeasures Dispensing System (Sheet 2 of 2)

Change 1

8-43

Section VIII

NAVAIR 01-H1AAB-1

CONTROL/INDICATOR

FUNCTION

1 .

MA indicator

Flashing indicates high radar missile threat with DSCRM switch is ON.

BRIL control

Adjusts indicator illumination.

NIGHT-DAY control

Adjust indicator intensity.

AUDIO control

Adjusts radar warning audio volume.

DSCRM switch:

OFF

Without missile activity - Provides strobe lines for ground radar and normal audio
indications.

With missile activity - Provides strobe lines for ground radar, flashing strobe line(s)
for missile activity, and flashing MA (missile alert) light.

Without missile activity — No indications.

With missile activity - Flashing strobe lines for missile activity (no strobe lines for
ground radar), flashing MA light, and audio warning.

SELF TEST switch:
with DSCRM switch OFF

PWR switch ON.

(NOTE: One minute warmup)
Monitor CRT and audio &
press and hold SELF TEST

Forward and aft strobes appear, extending to approximately the third circle on the
indicator graticule and 2.5 kHz PRF audio present immediately.

Rotate indicator BRIL
control CW & CCW

Within approximately 6 seconds, alarm audio present and MA lamp starts flashing.

Rotate control unit AUDIO
control between maximum
CCW and maximum CW

Indicator strobes brighten (CW) and dim as control is rotated.

Release SELF TEST

AUDIOS not audible at maximum CCW and clearly audible at maximum CW.

Set DSCRM to ON.

Press & hold SELF TEST

All indications cease.

Within approximately 4 seconds, a FWD or AFT strobe and 1.2 kHz PRF audio
present. Within approximately 6 seconds, the other strobe will appear and APRF
audio will double.

PWR switch:

Applies power to radar set.

OFF

De-energizes radar set.

N2/83

Figure 8-24. Radar Warning Indicator and Control AN/APR-39

8-44

Change 1

NAVAIR 01-H1AAB-1

Section VIII

Figure 8-25. AN/ALQ-144 IR Jammer System Control

N2/83

Figure 8-26. Countermeasure Equipment

Change 1

8-45/(8-46 blank)

NAVAIR 01-H1AAB-1

Section IX

SECTION IX — FLIGHT CREW
COORDINATION

TABLE OF CONTENTS

T 4 v* f \ 11 r» 4 1 /"\ y\

. . 9-1 Standard Terminology.

.9-2

.... 9-1 TOW Mission Coordination.

.9-2

XT /-v /^VAIITW^ AVVI rVOVC

9-1 Gunner Acquisition.

.9-2

Tactical Missions/

Pilot Acquisition.

.9-2

Training.

.9-2 TOW Launch.

.9-3

INTRODUCTION

While the AH-1T (TOW) can be flown single pilot,
the combat mission requires two pilots to occupy
the crew positions. A qualified observer or enlisted
non-crewmember may occupy the front cockpits
on some flights not requiring crew duties of that
person. Coordination between the two personnel
occupying the crew positions is absolutely neces¬
sary to enhance the mission capability and safety
of the crew.

Observer.

Any person, authorized by the commanding
officer or his designated deputy, may occupy the
front cockpit if the following requirements are
met:

1. Must complete an egress drill.

2. Must be fully briefed on the front cockpit.

3. Must have a current physical.

4. Must be fully briefed on what is expected of
him during the flight to include but not be
limited to:

a. Be alert for other aircraft or obstacles to
flight.

b. Operating altitudes.

c. Mission plan.

d. Actions during an emergency.

e. Lost communication with the pilot.

Non-crewmembers.

Non-crewmembers are designated in writing by the
commanding officer and are assigned to temporary-
definite orders involving flying. Those personnel
are the only personnel that shall fly, occupying
the front crew position in a non-crew status. In
addition to receiving the same information; prior
to flight, as outlined in the observer paragraph, he
will perform duties as directed by the aircraft
commander. Those duties may include but are not
limited to:

1. Assisting the aircraft commander in preparing
the aircraft for flight.

2. Acting as an observer.

3. Recording data as directed by the aircraft
commander.

Non-crewmembers shall meet the following require¬
ments and all others required by applicable
directives:

1. Must have a current flight physical.

2. Must be a 2nd class or better swimmer.

3. Must have current physiology training.

4. Must have current WST training.

These requirements are not to be interpreted as
limiting in any way the establishment of higher

9-1

Section IX

NAVAIR 01-H1AAB-1

requirements by proper authority. Non-crew¬
members ground training should include but not
be limited to:

1. Ground handling — Instructions in the opera¬
tion and use of all ground support equipment,
aircraft towing, and tiedown procedures (air¬
craft security). Instructions in the use of
proper taxi director signals, both day and
night.

2. Fueling and servicing — Instructions in the
proper fueling and servicing procedures with
particular emphasis on safety precautions,
fuel contaminations, alternate fuels, oils, and
lubricants.

3. Equipment stowage — Instructions in the
proper location and stowage of loose
equipment.

4. Helicopter inspection — Instructions in
assisting the aircraft commander in inspecting
the aircraft and securing aircraft panels,
doors, etc.

5. Fire guard — Instructions in procedures for
performing duties of fire guard during starts.

The fire guard will remain clear of the
tip path plane, engine compressor, and
engine turbine areas during start.

TACTICAL MISSIONS/TRAINING.

To enhance the mission capability of the AH-1T
(TOW), both pilots shall fully understand all duties
expected of him during the mission. By dividing
the duties required of crew, individual workloads
will be minimized. For example: The pilot that is
not controlling the aircraft should allow the other
pilot to copy all required communications. No
attempt is made here to cover all situations as the
complexity and variance of the AH-1 tasks is great.
However, the crew shall divide as many duties as is
practical, limited only by individual experiences
and proficiency at certain tasks.

Standard Terminology.

Much confusion can result in the cockpit due to
the non-standardized communication. Communica¬
tion with air traffic control agencies shall be

standardized in accordance with applicable direc¬
tives. Inner cockpit voice procedures should be
standardized by the units to expedite the
communication and rule out misunderstandings.
As an example: “I’ve got it”, a much used term,
should not be used. The following terminology
may be used and will rule out any confusion as to
the action being taken:

1. “I’ve got the traffic.”

2. “I’ve got the controls.”

3. “I’ve got the fuel switch.”

4. “I’ve got the brief.”

5. “I’ve got the obstacles.”

Communications shall be expeditious, clear,
concise, and understood by both crewmembers.

TOW Mission Coordination.

Due to the complexity of the TMS on the AH-1T
(TOW), cockpit workloads for both pilots have
increased over the standard AH-1T. When the co¬
pilot is utilizing the TSU, he is totally oriented on
that system and cannot aid the pilot in any other
task, such as, navigation, obstacle avoidance, or
communications. The following procedures are
recommended to expedite communication, target
acquisition, and employment of the TMS.

GUNNER ACQUISITION.

1. Gunner acquires target with GHS and states,
“GUNNER ACQ”.

2. Pilot states, “ROGER ACQ,” and attempts
to give the gunner as stable a platform as
possible to acquire the target. By noting the
gunner’s head position, the pilot can deter¬
mine the direction in which the aircraft must
be maneuvered to get into pre-launch
constraints.

3. When the gunner acquires the target through
the TSU, he states, “TARGET ACQ”.

PILOT ACQUISITION.

1. When the pilot has acquired a target with the
PHS, he states, “PILOT ACQ”.

9-2

NAVAIR 01-H1AAB-1

Section IX

2. The gunner acknowledges, “ROGER, PILOT
ACQ”, looks through the TSU and depresses
the PHS ACQ button. (The pilot may have to
describe the target).

3. When the gunner acquires the target he states,
“TARGET ACQ” and releases the PHS ACQ
button. Until the gunner states, “TARGET
ACQ”, the pilot must hold his PHS reticle
on the target.

TOW LAUNCH.

1. When the ATTK indicator appears on the PSI
the pilot states, “ATTACK”.

2. The gunner should direct the pilot to
maneuver the aircraft if an obstacle masks the
target, or if a more desirable line of flight can
be achieved to minimize aircraft exposure. He
should do this by stating, “COME UP, COME
DOWN, COME LEFT OR COME RIGHT.”
The pilot should also communicate to the
gunner the need for maneuvering the aircraft
to a more desirable firing position.

3. If wind conditions in a hover position neces¬
sitate momentarily rocking into constraints,
the pilot shall state, “ROCKING INTO
PITCH (OR ROLL).” If a steady constraints
is necessary, the pilot shall state, “DRIFTING
RIGHT (OR FORWARD) etc.”

4. When the pilot has an indication of system
armed, ATTACK flag, and he is ready for the
gunner to fire, he will state, “CLEARED TO
FIRE”. At this time, if the “ROCKING INTO
CONSTRAINTS” was given, the gunner
should hold the LHG TRIGGER down to
initiate the launch when constraints are
achieved.

5. After launch the pilot should keep the gunner
apprised as to the direction of aircraft
maneuvering. The gunner should roger unless
it appears to interfere with the TSU to target
line of sight. In that case, he should state,
“NEGATIVE, COME LEFT, RIGHT, etc.”

6. At wire cut the pilot who first notes it should
so state. If wire cut is caused by missile impact
the gunner should state, “IMPACT”. At
extended ranges, the pilot should not attempt
to discern whether the impact was the missile
or fire from another source.

9-3/(9-4 blank)

NAVAIR 01-H1AAB-1

Section X

SECTION X — NATOPS EVALUATION

TABLE OF CONTENTS

Concept.10-1

Implementation.10-1

Definitions .10-1

Ground Evaluation.10-2

Grading Instructions.10-2

Flight Evaluation.10-3

Flight Evaluation Grading
Criteria.10-4

CONCEPT.

The standard operating procedures prescribed in this
manual represent the optimum method of operating
AH-1T aircraft. The NATOPS Evaluation is
intended to evaluate compliance with NATOPS
procedures by observing and grading individuals and
units. This evaluation is tailored for compatibility
with various operational commitments and missions
of both Navy and Marine Corps units. The prime
objective of the NATOPS Evaluation program is to
assist the unit commanding officer in improving unit
readiness and safety through constructive comment.
Maximum benefit from the NATOPS Program is
achieved only through the active vigorous support of
all pilots and flight crewmembers.

IMPLEMENTATION.

The NATOPS Evaluation program shall be carried
out in every unit operating naval aircraft. The
various categories of flight crewmembers desiring to
attain/retain qualification in the AH-IT shall be
evaluated in accordance with OPNAV Instruction
3510.9 series. Individual and unit NATOPS
Evaluations will be conducted periodically; however,
instructions in and observation of adherence to
NATOPS procedures must be on a daily basis within
each unit to obtain maximum benefits from the
program. The NATOPS coordinators, Evaluators,
and Instructors shall administer the program as
outlined in OPNAVINST 3510.9 series. Evaluees
who receive a grade of Unqualified on a ground or
flight evaluation shall be allowed 30 days in which to
complete a re-evaluation. A maximum of 60 days
may elapse between the date the initial ground
evaluation was commenced and the date the flight
evaluation is satisfactorily completed.

Final Grade Determination.10-5

Records and Reports.10-5

NATOPS Evaluation Report (OPNAV Form

3510-8).10-6

AH-1T (TOW) NATOPS Open Book Exam . . . 10-7

DEFINITIONS.

The following terms, used throughout this section,
are defined as to their specific meaning within the
NATOPS program.

NATOPS Evaluation.

A periodic evaluation of individual flight
crewmember standardization consisting of an open
book examination, a closed book examination, an
oral examination, and a flight evaluation.

NATOPS Re-evaluation.

A partial NATOPS Evaluation administered to a
flight crewmember who has been placed in an
Unqualified status by receiving an Unqualified
grade for any of his ground examinations or the flight
evaluation. Only those areas in which an
unsatisfactory level was noted need be observed
during a re-evaluation.

Qualified.

That degree of standardization demonstrated by a
very reliable flight crewmember who has a good
knowledge of standard operating procedures and a
thorough understanding of aircraft capabilities and
limitations.

Conditionally Qualified.

That degree of standardization demonstrated by a
flight crewmember who meets the minimum
acceptable standards. He is considered safe enough to
fly as a pilot in command or to perform normal duties

10-1

Section X

NAVAIR 01 -HI AAB-1

without supervision but more practice is needed to
become Qualified.

Unqualified.

That degree of standardization demonstrated by a
flight crewmember who fails to meet minimum
acceptable criteria. He should receive supervised
instruction until he has achieved a grade of Qualified
or Conditionally Qualified.

Area.

A routine of preflight, flight, or postflight.

Sub-Area.

A performance sub-division within an area, which is
observed and evaluated.

Critical Area/Sub-Area.

Any area or sub-area which covers items of
significant importance to the over-all mission
requirements, the marginal performance of which
would jeopardize safe conduct of the flight.

Emergency.

An aircraft component, system failure, or condition
which requires instantaneous recognition, analysis,
and proper action.

Malfunction.

An aircraft component or system failure or condition
which requires recognition and analysis, but which
permits more deliberate action than that required for
an emergency.

GROUND EVALUATION.

Prior to commencing the flight evaluation, an
evaluee must achieve a minimum grade of Qualified
on the open book and closed book examinations. The
oral examination is also part of the ground
evaluation but may be conducted as part of the flight
evaluation. To assure a degree of standardization
between units, the NATOPS Instructors may use the
bank of questions contained in this section in
preparing portions of the written examinations.

Open Book Examination.

The open book examination may consist of but shall
not be limited to the questions from the question
bank. The number of questions shall not exceed that
of the question bank nor be less than 50. The purpose
of the open book examination portion of the written
examination is to evaluate the crewmembers
knowledge of appropriate publications and the
aircraft. The maximum time for this examination
should not exceed seven days.

Closed Book Examination.

The closed book examination may consist of but shall
not be limited to the questions from the question
bank. The number of questions on the examination
will not exceed 40 or be less than 20. Questions
designated critical will be so marked. An incorrect
answer to any question in the critical category will
result in a grade of unqualified being assigned to the
examination.

Oral Examination.

The questions may be taken from this manual and
drawn from the experience of the
Instructor/evaluator. Such questions should be direct
and positive and should in no way be opinionated.

OFT/WST Procedures Evaluation (If Applicable).

An OFT may be used to assist in measuring the
crewmembers efficiency in the execution of normal
operating procedures and his reaction to emergencies
and malfunctions. In areas not served by these
faci 1 ities, this may be done by placing the
crewmember in an aircraft and administering
appropriate questions.

GRADING INSTRUCTIONS.

Examination grades shall be computed on a 4.0 scale
and converted to an adjective grade of Qualified or
Unqualified.

Open Book Examination.

To obtain a grade of Qualified, an evaluee must
obtain a minimum score of 3.5.

Closed Book Examination.

To obtain a grade of Qualified, an evaluee must
obtain a minimum score of 3.3.

10-2

NAVAIR 01-H1AAB-1

Section X

Oral Examination and Oft Procedure Check.

AIR TAXI.

(If conducted.)

1. Taxi.

A grade of Qualified or Unqualified shall be assigned
by the Instructor/Evaluator.

TAKEOFF/TRANSITION.

1. Procedures.

FLIGHT EVALUATION.

The NATOPS flight evaluation is intended to
evaluate unit/individual compliance with approved
standardized operating procedures. The successful
completion of all ground evaluations and
examinations is required prior to commencement of
the flight evaluation. Insofar as possible, evaluation
flights will be scheduled so as not to interfere with
squadron operations. The flight evaluation should
conform to any syllabus flight. Only those areas
observed or required by the mission will be
evaluated. Determination of the final flight
evaluation grade will be made as outlined in the
Final Grade Determination section.

2. Type takeoff.

a. Vertical.

b. Cross-wind.

c. Maximum gross.

3. Transition.

CLIMB/CRUISE.

1. Procedures.

2. Power control.

NOTE

Areas/sub-areas to be evaluated are listed.

Critical areas/sub-areas are marked by an
asterisk.

3. Helicopter control.

APPROACH AND LANDING

Pilot's Nontactical Flight Evaluation.

1. Procedures.

MISSION PLANNING.

2. Power control.

1. Flight plan.

3. Helicopter control.

2. Computation card.

4. Type of landing:

3. Weather.

a. Vertical.

BRIEFING.

b. Running.

PREFLIGHT.

c. Cross-wind.

1. Records check.

D. Maximum gross.

2. Preflight check.

AUTOROTATION*.

3. Crew briefing.

1. Procedures.

ENGINE AND ROTOR START.

2. RPM control.

1. Start.

3. Airspeed control.

2. Post start.

4. Recovery.

10-3

Section X

NAVAIR 01-H1AAB-1

EMERGENCY PROCEDURES'

1. Procedures.

2. Helicopter control.

CREW COORDINATION.

DEBRIEFING.

MISSION FLIGHT EVALUATION.

CONFINED AREA LANDING PRECISION
APPROACH.

1. Procedures.

2. Approach.

3. Power control.

4. Helicopter control.

NAVIGATION.

SEARCH AND RESCUE.

SPECIAL.

Crewmember — Evaluation areas.

1. Preflight.

2. Security.

3. Ground safety precautions.

4. Hand signals.

5. Fueling and servicing of aircraft.

6. Post flight.

7. Emergency Procedures*.

8. Rescue operations (coordination and cover).

FLIGHT EVALUATION GRADING
CRITERIA.

Only those sub-areas provided or required will be
graded. The grades assigned for a sub-area shall be
determined by comparing the degree of adherence to
standard operating procedures with adjectival

ratings listed below. Momentary deviations from
standard operating procedures should not be
considered as unqualifiying provided such deviations
do not jeopardize flight safety and the evaluee applies
prompt corrective action.

Qualified.

Well standardized evaluee demonstrated highly
professional knowledge of and compliance with
NATOPS standards and procedures; momentary
deviations from or minor omissions in non-critical
areas are permitted if prompt at timely remedial
action is initiated by the evaluee.

Conditionally Qualified.

Satisfactorily standardized; one or more significant
deviations from NATOPS standards and procedures,
hut no errors in critical areas and no errors jeopardize
mission accomplishment of flight safety.

Unqualified.

Not acceptably standardized; evaluee fails to meet
minimum standards regarding knowledge of and/or
ability to apply NATOPS procedures; one or more
significant deviations from NATOPS standards and
procedures which could jeopardize mission
accomplishment or flight safety.

Flight Evaluation Grade Determination.

The following procedure shall be used in determining
the flight evaluation grade: A grade of Unqualified in
any critical area/sub-area will result in an overall
grade of Unqualified for the flight. Otherwise, flight
evaluation (or area) grades shall be determined by
assigning the following numerical equivalents to the
adjective grade for each sub-area. Only the numerals
0, 2, or 4, will be assigned in sub-areas. No
interpolation is allowed.

Unqualified.0.0

Conditionally Qualified .2.0

Qualified .4.0

To determine the numerical grade for each area and
the overall grade for the flight, add all the points
assigned to the sub-areas and divide this sum by the

10-4

NAVAIR 01 -HIAAB-1

Section X

number of sub-areas graded. The adjective grade 3.0 to 4.0.Qualified

shall then be determined on the basis of the following

cfalp EXAMPLE: (Add Sub-area numerical equivalents)

0.0 to 2.19 . ..

.Unqualified

4 + 2 + 4 + 2 + 4=16 = 3.20 Qualified

rr r

2.2 to 2.99 . ..

, . . .Conditionally Qualified

5 o

FINAL GRADE DETERMINATION

The final NATOPS Evaluation grade shall be the same as the grade assigned to the flight evaluation. An
evaluee who receives an Unqualified on any ground examination or the flight evaluation shall be placed in an
Unqualified status until he achieves a grade of Conditionally Qualified or Qualified on a re-evaluation.

RECORDS AND REPORTS

A NATOPS Evaluation report (OPNAV Form 3510-8) shall be completed for each evaluation and forwarded to
the evaluee’s commanding officer. Refer to figure 10-1.

This report shall be filed in the individual flight training record and retained therein for 18 months. In addition,
an entry shall be made in the pilot/NFO flight log book under Qualifications and Achievements as follows.

QUALIFICATION

DATE SIGNATURE

NATOPS

EVAL.

(Aircraft

(Crew

(Date)

(Authenticating

(Unit which

Model)

Position)

Signature)

Administered

Eval.)

10-5

Section X

NAVAIR 01-H1AAB-1

NA TOPS EVALUA TION REPORT

(OPNA V FORM 3510-8)

NATOPS EVALUATION REPORT

OPNAV FORM 3510/8 (REV. 10-73) S/N 0107-LF-723-0001

REPORT SYMBOL OPNAV 3510-3

NAME (Last, First Initial)

GRADE

SSN

SQUADRON/UNIT

AIRCRAFT MODEL

CREW POSITION

TOTAL PILOT/FLIGHT HOURS

TOTAL HOURS IN MODEL

DATE OF LAST EVALUATION

NATOPS EVALUATION

□

CHECK IF CONTINUED ON REVERSE SIDE

GRADE, NAME OF EVALUATOR/INSTRUCTOR

SIGNATURE

DATE

GRADE, NAME OF EVALUEE

SIGNATURE

DATE

REMARKS OF UNIT COMMANDER

RANK, NAME OF UNIT COMMANDER

SIGNATURE

DATE

*WST, OFT, COT, or cockpit check in accordance with OPNAVINST 3510.9E

«■ 002147

10-6

Figure 10-1. NATOPS Evaluation Report (OPNAV Form 3510/8)

NAVAIR 01-H1AAB-1

Section X

Name__

Date__

Score __

AH-1T TOW NATOPS OPEN BOOK EXAM

1. Should conflict exist between the NATOPS Flight Manual and other publications, the NATOPS
Flight Manual shall govern. TRUE/FALSE.

2. Anyone can recommend a change to NATOPS. TRUE/FALSE.

3. A WARNING as defined in NATOPS, as an operating procedure or technique which

may----

4. The T400-WV-402 engine is a twin power section turboshaft engine consisting of_

sections driving a single output shaft through separate halves of a common __.

5. Max.gross weight of the AH-1T (TOW) is_pounds.

6. The AH-1T has a rotor diameter of-feet and an overall length with the main rotor in the fore

and aft position, and the tail rotor in the horizontal position of-feet.

7. The compressor section of the T400-WV-402 engine contains-axial and-centrifugal

stages.

8. Fuel is sprayed into the annular combustion change by-fuel nozzles.

9. The_ oil cooler system has an automatic emergency oil cooler bypass valve that routes the

oil around the oil cooler or lines, if the oil cooler or lines are ruptured.

10. There are-independent oil systems within the power plant.

11. A-second delay is built into the engine idle stop release switch to allow time to-or-

throttle.

12. The torquemeter system receives power from the-bus and is protected by the-

circuit breaker.

13. The triple tachometer is powered by the 115 vac essential bus. TRUE/FALSE.

14. During a complete electrical failure there will be no indication of rotor RPM. TRUE/FALSE.

15. The gas producer turbine tachometers operate independently of the electrical system.

TRUE/FALSE.

16. The combining gearbox oil system does not incorporate an oil hot caution light. TRUE/FALSE.

17. The transmission oil bypass valve closes automatically because of-flow between the pump

and-outlet.

18. The fuel boost pumps are powered by the-bus-.

10-7

Section X

NAVAIR 01-H1AAB-1

19.

20 .

21 .

22 .

23.

24.

25.

26.

27.

28.

29.

30.

31.

32.

33.

34.

35.

36.

Movement of either fuel switch to ON energizes both fuel boost pumps. TRUE/FALSE.

With the crossfeed valve closed, the forward cell supplies engine_only and the aft cell supplies

engine_only.

The fuel pressure indicator reads-pump pressure.

In the event of a fuel filter caution light, if clogging continues, the_opens to allow

fuel to-the clogged filter.

With the FUEL TANK INTCON switch in the CLOSE position, the AFT FUEL LOW light
illuminates when .-pounds of fuel remains.

One air pump can pressurize both auxiliary fuel tanks if a failure of one pump occurs. TRUE/FALSE.
Power for turret control and firing is supplied by the No_generator.

In the event of failure of the No. 2 generator while supplying turret power, the _. will

automatically switch to supply turret power. The main bus is then supplied by the__

The primary electrical power supply system is a-single-wire, negative-ground_

arrangement supplied by two-, 200 ampere-, one mounted on each engine.

When the NON-ESS BUS switch is in NORMAL, power is supplied to the non-essential bus as long
as-is operating.

In the event of a MAIN inverter failure, use of the TACAN can be regained by placing the NON-ESS
BUS switch in MANUAL. TRUE/FALSE.

The external power receptacle incorporates overvoltage protection. TRUE/FALSE.

The hydraulic pumps deliver-psi output pressure at___rpm.

Hydraulic system No. 1 supplies system power for the_,_ > and _

actuators and the_SCAS actuator.

The rotor brake is powered by the No._hydraulic system.

The hydraulic filter indicator will pop out when the differential pressure across the filter element
exceeds-psi.

The SCAS channel engage switches energize electric__ valves controlling __

to the system.

The pointer located at the 6 o’clock position on the pilot’s attitude indicator will deviate toward an
FM station when the FM control panel mode selector switch is in the HOME position and a usable
signal is received. TRUE/FALSE.

10-8

NAVAIR 01-H1AAB-1

Section X

37. The pilot controls the RESET functions of the copilot/gunner_

light. He also controls the BRIGHT/DIM function of the copilot/gunner_lights.

38. Do not actuate the FIRE WARNING TEST switch more than_seconds.

39. Pulling the FIRE PULL handle will shut off-to the affected engine, deactivate the_.

and-circuits, close the _ , and_both fire extinguisher bottles.

40. If both FIRE PULL HANDLES are pulled out and the FIRE EXT switch is moved to_

position, the bottle-discharge.

41. To actuate the canopy jettison system, rotate any handle_degrees and_

42. The inertia reel will automatically lock the shoulder harness when the helicopter encounters

an impact force in excess of-G deceleration.

43. For weight and balance purposes, the AH-1T (TOW) is classified as a class-helicopter.

44. The ECU and RAIN RMV switches shall be off for takeoff, landing or any time_

is required.

45. The fuselage formation lights are powered by the-bus.

46. At temperatures of -25 degrees C and below, engine oil should be changed to MIL-L--

47. Total fuel capacity is-U.S. gallons, of which -gallons are unusable.

48. The forward ground handling gear should be used when the helicopter is at a-

and/or-of mid CG.

49. The duty cycle for the starter is:-seconds on,-minute(s) off,-seconds on, -minute(s)

off,-seconds on,-minute(s) off.

50. The transmission torque limit in a dive is-%.

51. Power off, maximum transient rotor RPM is_ %.

52. The maximum continuous ITT limit is-degrees C.

53. For engine starting, the 5 second limit is-to-degrees C. The two second limit is-to

_degrees C.

54. Without stores, maximum airspeed is_KIAS.

55. In a steady state autorotation, maximum airspeed is_KIAS.

56. The airspeed indicator is unreliable below-KIAS.

57. With wing stores, maximum airspeed is-KIAS.

58. Decrease airspeed-KIAS for each 1,000 feet of density altitude above-feet.

59. No airstarts or manual fuel control operation are permitted above_feet.

Change 1

10-9

Section X

NAVAIR 01-H1AAB-1

The most critical flight regime with the lateral CG at the most left station is a __KIAS__

At a gross weight of 12,500 pounds, the aft CG limit is fuselage station__

The effect of humidity on gas turbine engines is negligible. TRUE/FALSE.

If a starting attempt is discontinued, allow the engine to come to a _ and then

accomplish a_second_run.

A non-engaged engine is indicated by- slightly higher than the engaged engine and a near

zero-indication.

The rain removal system shall not be utilized on a_windshield.

After takeoff, takeoff power should be maintained until a safe_airspeed is attained.

When making a slope landing, if mast bumping occurs, reposition_toward__

Rapid application of-at or near flat pitch can result in a _or_overspeed.

Reducing collective rapidly and applying_cyclic can result in_overspeed.

Without the use of rotor brake on shutdown, winds of approximately_knots or above may

cause the rotor to windmill indefinitely.

Carrier qualifications remain current for_months.

At gross weights of-pounds or lower,the aft ground handling gear may be used for moving

the helicopter.

Aboard ship, in an emergency, the helicopter may be launched in_knot relative winds.

Aboard ship at night, the 180 degree position is-yards abeam.

Aboard ship, a-from the helicopter director is mandatory.

Full autorotative landings may be practiced by pilots —-by_authority.

During autorotation, if the helicopter is only slightly out of balanced flight, the rate of descent will
be increased by about -feet per minute.

At average gross weights, best glide speed is approximately_knots.

The two basic types of formations are- and __ .

The parade position for echelon, fingertip and diamond is on a_degree bearing either side

of lead axis with -feet of step up.

A marked increase in airframe vibration and, possibly, control feedback is an indication of impending

10-10

NAVAIR 01-H1AAB-1

Section X

82.

83.

84.

85.

86 .

87.

88 .

89.

90.

91.

92.

93.

94.

95.

96.

97.

98.

99.

During left rolling maneuvers or high power dives,-— , -, and-increases occur.

AH-1T (TOW) helicopters have a tendency to roll to the-when forward cyclic is used to

initiate a lower than_G maneuver in forward flight.

Mast bumping generally occurs at the-of the operating- .

During autorotation transient rotor rpms of-to-% are allowed up to-seconds.

List the four factors which affect power-off rotor rpm.

1 .

2 .

If an engine fails during takeoff, -- , -,-* and -

will determine if flight can be maintained.

With a dual hydraulic failure, cyclic feedback may be encountered at airspeeds below-KIAS.

A shear pin is incorporated in the --linkage connection to the collective

linkage.

For most gross weights, it is unlikely that the AH-1T (TOW) can achieve a---

flig ht condition following loss of tail rotor thrust.

In an ACTUAL emergency, it is not necessary to wait for Ng to stabilize at engine idle before
switching to manual fuel. TRUE/FALSE.

Loss of DC power from the-—-—-to the main inverter will result in

_ switch over to the standby inverter.

Total loss of electrical power will cause the loss of all engine and component instruments, indicators,
and gages except---tachometers.

The canopy doors may be opened in flight below-KIAS.

Avoid helicopter operation with dual fuel boost pump failure above-feet.

Under certain conditions, airspeed in excess of-KIAS may be necessary to land under single

engine conditions.

Do not operate the engine in excess of-% Ng until engine and combining gearbox oil tempera¬
ture reach +-degrees C.

When the microphone at the right crew station is keyed, it connects the right station to
the_mike circuit.

When testing the radar altimeter, a reading of
if the system is functioning properly.

plus or minus_feet will be indicated

Change 1

10-11

Section X

NAVAIR 01-H1AAB-1

The azimuth range of the turret system is-degrees, the gun may be depressed a maximum

ot-degrees.

If the mode switch on the NARCADS panel is placed in ALL, the_

function is disabled.

The smoke grenade system is energized with the master arm switch in STBY. TRUE/FALSE.

The PILOT OVERRIDE switch electrically bypasses the______

The gunner ACQ switch is located on the____

The TOW missile programmer programs the firing sequence for a period of

-- TOW missiles may be carried on the _

The TOW missile launch motor is expended before the

wing stations.

The TSU has an angular field of view of_degrees in low mag and

The ready flag on the PSI means _____

The Weapon Control switch has to be in ___

Reinstallation of the upper TMLs will require that the system be_

degrees in high mag.

to operate the TMS.
- before use.

When in PILOT OVERRIDE mode, the gunner utilizes the
switches to direct and fire the turret utilizing the GHS.

10-12

NAVAIR 01-H1AAB-1

Section XI
Part 1

SECTION XI — PERFORMANCE DATA

TABLE OF CONTENTS

Introduction.11-1 PART 4 — RANGE

PART 1 - STANDARD DATA Best ran ^ e .1M8

Range.11-18

Airspeed calibration .11-1 Time and ran S e versus fuel.11-18

Density altitude.11-2

Shaft horsepower.11-2

Torque available.11-2

Fuel flow .11-2

PART 2 - TAKEOFF

Maximum gross weight for hovering .... 11-10
Indicated torque required to hover.11-10

PART 3 - CLIMB

Climb performance.11-15

Service ceiling.11-15

INTRODUCTION.

The charts presented on the following pages are
provided to aid in preflight and in-flight planning.
Through the use of the charts, the pilot is able to
select the best power setting, altitude, and airspeed

PART 5 — ENDURANCE

Maximum endurance.11-49

Hover endurance.11-49

PART 6 - EMERGENCY OPERATION

Ability to maintain flight on one

engine.11-55

Minimum airspeed for flight with one
engine.11-55

PART 7 - SPECIAL CHARTS
Radius of turn at constant airspeed.11-87

to be used to obtain optimum performance for the
mission being flown. The charts are presented in
graphic or profile form. Charts are based on flight-
test data, estimated data or calculated data as
indicated on the chart.

PART 1 — STANDARD DATA

AIRSPEED CALIBRATION CHART. PRESSURE ALTITUDE.

The airspeed calibration chart (figure 11-1)
converts calibrated airspeed to indicated airspeed
and vice versa. Calibrated airspeed (KCAS) is
indicated airspeed (KIAS) as read from the
airspeed indicator corrected for instrument error,
plus the installation correction. Corrections for
cruise, climb, and autorotation are shown.

EXAMPLE: Convert 131 KCAS to equivalent
KIAS for a cruise condition.

Solution:

a. Enter figure 11-1 at 131 KCAS. Move
right to cruise line.

b. Drop down and read 136.8 KIAS.

Pressure altitude is the altitude indicated on the
altimeter when the barometric scale is set on 29.92.
It is the height above the theoretical plane at
which the air pressure is equal to 29.92 inches of
mercury.

DENSITY ALTITUDE.

Density altitude is an expression of the density of
the air in terms of height above sea level; hence,
the less dense the air, the higher the density
altitude. For standard conditions of temperature
and pressure, density altitude is the same as
pressure altitude. As temperature increases above
standard for any altitude, the density altitude will

11-1

Section XI
Part 1

NAVAIR 01 -HIAAB-1

also increase to values higher than pressure
altitude. Figure 11-2 expresses density altitude as a
function of pressure altitude and temperature.

The chart also includes the inverse of the square
root of the density ratio (1 /yfo~ ), which is used to
calculate TAS by the relation:

TAS = CAS x 1/Vo"

EXAMPLE: If the ambient temperature is 0°C
(standard day) and the pressure altitude is 4000
feet, find the density altitude, l/>/o", and true
airspeed for 131 KCAS.

Solution:

a. Enter the bottom of the chart (figure 11-2)
at +0°C.

b. Move vertically upward to the 4000 feet
pressure altitude line.

c. From this point, move horizontally to the
left and read a density altitude of 3150
feet and move horizontally to the right
and read l/y/o' equals 1.047.

d. True airspeed = KCAS x 1 /\fo =

131 x 1.047 - 137.2 KTAS

SHAFT HORSEPOWER VERSUS
TORQUE.

The shaft horsepower versus torque chart (figure
11-3) provides a means of converting torque to
shaft horsepower, and vice versa, for 100 percent
rotor rpm.

EXAMPLE: Determine the shaft horsepower
equivalent for a 37 percent torque (%Q) during
single engine operation, 100 percent rotor rpm.

Solution:

a. Enter figure 11-3 at 37%Q single engine
torque. Move up and intersect the
baseline.

b. Move left, read 750 single engine shaft
horsepower.

TORQUE AVAILABLE.

Outside air temperature and pressure altitude
change the capability of the turboshaft engine to
produce power at the rated interturbine
temperatures. Figures 11-4 and 11-5 shows the
power available at the intermediate and the
maximum continuous power ratings respectively
for both twin and single engine operation.

The torque output capability of the engine can
exceed the structural limits of the transmision
under certain conditions during twin engine
operation. Because of this, the restriction on twin
engine operation is (transmission torque) 100%Q
for 5 minutes operation, and 84.9%Q for
continuous operation. The restrictions on single
engine operation are due to engine mechanical
limitations, therefore 53.1%Q for 30 minutes and
45.2%Q for continuous operation shall not be
exceeded during single engine operation.

EXAMPLE: At 4000 feet, +7.1°C (standard day)
find the maximum continuous power available
during both twin and single engine operation.

Solution:

a. Figure 11-5 shows maximum continuous
power available. Enter the left scale at
4000 feet pressure altitude.

b. Move right and interpolate for +7.1°C
OAT.

c. For twin engine torque move up and read
70%Q twin engine torque available, or for
single engine torque move down and read
35%Q single engine torque available.

FUEL FLOW.

The fuel flow chart (figure 11-6) shows the fuel flow
for a given altitude and power setting for both twin
and single engine operation at 0°C outside air
temperature.

EXAMPLE: Find the fuel consumption of the
helicopter at a torquemeter setting of 55%Q for
4000 feet, 0°C OAT, twin engine operation.

11-2

NAVAIR 01-H1AAB-1

Section XI
Part 1

Solution:

a. Enter the top torque scale of the chart
(figure 11-6) at 55 percent indicated
torque.

b. Move down to the 4000 feet pressure
altitude line.

c. From this intersection, move right and
read a twin engine fuel flow of 750 pounds
per hour.

11-3

Section XI NAVAIR 01-HI AAB-1

Part 1

11-4

NAVAiR 01 -HIAAB-1

Section XI
Part 1

Figure 11-2. Density altitude/temperature conversion chart

DENSITY ALTITUDE/TEMPERATURE CONVERSION

DATA BASIS: CALCULATED DATA

OAT — ° F

OAT — °C

.36

.20

.16

Vo-

.08

.04

.00

11-5

Section XI
Part 1

NAVAIR 01 -HIAAB-1

£5MIN

XMSN

LIM

Figure 11-3. Shaft horsepower vs torque chart

SHAFT HORSEPOWER VS TORQUE

100% ENGINE RPM

MODEL: AH-IT(TOW)

DATE: 1 AUGUST 1978
DATA BASIS: FLIGHT TEST

ENGINE: T400-WV-402
FUEL GRADE: JP-4/JP-5
FUEL DENSITY: 6.5/6.8 LB/GAL

1 100 -

1000 -

900—

800—

700-

{/) 600-

400—

300-

200 —

100 -

qh ^^ i k i . ...

30 MIN ENG MECH LIM

TORQUE — % Q (TWO ENGINES)
40 50 60 70

15 20 25 30 35

TORQUE — % Q (SINGLE ENGINE)

-2200

100 110

o- 1

in*

2000

CONT ENG MECH LIM

1800

CjL /./ ( Af < ( djL

CONT
XMSN LIM

1600

1400 CD

1200

1000

600

400

200

11-6

NAVAIR 01-H1AAB-1

Section XI
Part 1

Figure 11-4. Maximum torque available (30 minute operation) chart

11-7

Section XI
Part 1

NAVAIR 01 -HIAAB-1

11-8

NAVAIR 01 -HIAAB-1

Section XI
Part 1

FUEL FLOW VS TORQUE

OAT = 0°C

100% ENGINE RPM

MODEL: AH-IT(TOW)

DATE: 1 AUGUST 1978
DATA BASIS: FLIGHT TEST

ENGINE T400-WV-402

FUEL GRADE JP-4/JP-5

FUEL DENSITY: 6 5/6 8 LB/GAL

TORQUE — % Q (TWO ENGINES)

LLI

Figure 11 -6. Fuel flow vs torque chart

Section XI
Part 2

NAVAIR 01 -HIAAB-1

PART 2 —

MAXIMUM GROSS WEIGHT FOR
HOVERING.

The maximum gross weight for hovering charts
(figure 11-7, sheets 1 and 2) and (figure 11-17,
sheets 1 and 2) presents data for twin engine and
single engine operation respectively. The charts
show the maximum gross weight hover capability
at a pressure altitude, outside air temperature
combination while at maximum torque available.
Effect of skid height above ground is shown in
each sheet 1, and effect of headwind is presented in
each sheet 2; of figures 11-7 and 11-17.

EXAMPLE: Find the maximum gross weight to
hover out of ground effect with zero and 10 knots
headwind at +7.1 C (standard day) and 4000 feet,
during twin engine operation.

Solution:

a. Enter figure 11-7, (sheet 1 of 2) on the left
at 4000 feet pressure altitude.

b. Move to the right to the twin engine
operation region. Interpolate between the
+5°C and +10°C lines for +7.1°C.

c. Drop down to the baseline, then read
12,700 pound maximum gross weight to
hover out of ground effect at 4000 feet,
standard day.

d. Enter figure 11-7 (sheet 2 of 2) at 12,700
pounds.

e. Move up to the baseline, following the
trend of the guidelines and move down to
the 10 knot headwind line.

f. Drop down, read 13,320 pound maximum
gross weight to hover OGE at 4000 feet,
standard day with a 10 knot headwind.

TAKEOFF

INDICATED TORQUE REQUIRED TO
HOVER.

The indicated torque required to hover charts
(figure 11-8, sheets 1 and 2) presents the torque
required to hover for various gross weights at
pressure altitudes between sea level and 12,000 feet
and outside air temperatures between plus or
minus 50°C. The effect of skid height is shown on
sheet 1 and the effect of headwind is presented on
sheet 2.

EXAMPLE: Find the & torque required to hover
OGE a 12,700 pound helicopter at 4000 feet,
standard day with a zero headwind condition and
with a 10 knot headwind condition.

Solution:

a. Enter figure 11-8, sheet 2 at 12,700 pound
gross weight.

b. Move right to 4000 feet pressure altitude.

c. Drop down to OAT baseline and follow
trend of the guidelines to +7.1°C
(standard day), from this intersection
drop down to the headwind baseline, read
81.5%Q OGE torque required to hover
(zero wind).

d. Move down, following trend of the
guidelines, to 10 knots headwind, from
this intersection drop down to torque
scale and read 75.5%Q OGE torque
required to hover (10 knot headwind).

e. In order to find the effect of headwind
torque required, subtract 10 knot
headwind condition from the zero wind
condition (81.5 - 75.5) = 6%Q & torque.

11-10

NAVAIR 01 -HI AAB-1

Section XI
Part 2

11-11

Section XI
Part 2

NAVAIR 01 -HIAAB-1

Figure 11-7. Maximum gross weight for hovering chart (Sheet 2 of 2)

11-12

NAVAIR 01 -HI AAB-1

Section XI
Part 2

INDICATED TORQUE REQUIRED TO HOVER

(EFFECT OF SKID HEIGHT ABOVE GROUND)

ZERO WIND CONDITION

MODEL: AH-IT(TOW) ENGINE: T400-WV-402

DATE: 1 AUGUST 1978 FUEL GRADE: JP-4/JP-5

DATA BASIS: FLIGHT TEST FUEL DENSITY 6 5/6 8 LB/GAL

TORQUE — % Q

Figure 11-8. Indicated torque required to hover chart (Sheet 1 of 2)

11-13

Section XI
Part 2

NAVAIR 01 -HI AAB-1

INDICATED TORQUE REQUIRED TO HOVER

(EFFECT OF HEADWIND)

OUT OF GROUND EFFECT

MODEL: AH-IT(TOW) ENGINE: T400-WV-402

DATE: 1 AUGUST 1978 FUEL GRADE: JP-4/JP-5

DATA BASIS: FLIGHTTEST FUEL DENSITY: 6.5/6 8 LB/GAL

TORQUE — % Q

Figure 11-8. Indicated torque required to hover chart (Sheet 2 of 2)

11-14

NAVAIR 01-H1AAB-1

Section XS
Part 3

PART 3

CLIMB PERFORMANCE.

The climb performance charts (figure 11-9, twin
engine operation; and figure 11-18, single engine
operation) show the time, distance, and fuel to
climb from sea level to 12,000 feet. Thus, when
climbing from any one altitude to another, the
climb performance is the difference between a
climb from sea level to the initial altitude, and a
climb from sea level to the final altitude.

These charts do not include the fuel used for
warmup and takeoff, which is 26 pounds (two
minutes at maximum continuous power). This
amount must be added to the climb fuel to
determine the total fuel required to reach cruise
altitude from an engine-off pretakeoff condition.

EXAMPLE: Find the time, distance, and fuel
required to climb from 4000 feet, +7.1°C (standard
day) to 10,000 feet, -4.8°C (standard day), with a
gross weight of 12,700 pounds, twin engine
operation.

Solution:

a. Enter the gross weight scale of figure 11-9
at 12,700 pounds. Proceed horizontally to
the right and intersect the 4000 feet
pressure-altitude curve. Move vertically
upward and interpolate for +7.1°C OAT.
Move horizontally to the time scale and
read 1.5 minutes.

b. Enter the gross weight scale of figure 11-9
again at 12,700 pounds. Proceed
horizontally to the right and intersect the
10,000 feet pressure altitude curve.
Move vertically upward and interpolate
for -4.8°C OAT. Move horizontally to the
time scale and read 5.9 minutes.

c. To obtain the time to climb from 4000 feet
to 10,000 feet, subtract the time to climb

— CLIMB

from sea level to 4000 feet from the time to
climb from sea level to 10,000 feet. For
this example, the time to climb would be
(5.9 - 1.5) = 4.4 minutes.

d. Using the same procedure as above,
distance to climb from 4000 to 10,000 feet
would be (7.5 - 1.2) = 6.3 nautical miles;
climb fuel from 4000 to 10,000 feet would
be (95 - 30) = 65 pounds.

e. If the climb began with a warmup and
takeoff, the climb fuel would include this
fuel allowance, or (65 + 26) = 91 pounds.

SERVICE CEILING.

There are two service ceiling charts shown. One,
figure 11-23, is for single engine operation at
intermediate rated power; the other, figure 11-10, is
for twin engine operation at maximum continuous
power. The single engine service ceiling chart is for
emergency situations where one engine is
inoperative and for planning purposes wherein the
pilot wishes to pick a route that does not rely on
both engines operating continuously.

EXAMPLE: Find the -5°C service ceiling for a
13,000 pounds gross weight helicopter at twin
engine operation, maximum continuous power.

Solution:

a. Since the helicopter is operating at twin
engine maximum continuous power, turn
to figure 11-10.

b. Enter figure 11-10 at 13,000 pounds gross
weight. Move straight upward and
intersect the -5°C line.

c. Move horizontally to the left and read a
service ceiling of 12,000 feet.

11-15

Section XI
Part 3

NAVAIR 01 -H1AAB-1

Figure 11-9. Climb performance chart

CLIMB PERFORMANCE

TWO ENGINE OPERATION AT INTERMEDIATE RATED POWER
ALL CONFIGURATIONS

MODEL: AH-IT(TOW)

DATE: 1 AUGUST 1978
DATA BASIS: FLIGHT TEST

100% ENGINE RPM
65 KIAS

ENGINE: T400-WV-402

FUEL GRADE JP-4/JP-5

FUEL DENSITY 6 5/6 8 LB/GAL

11-16

NAVAIR 01 -HIAAB-1

Section XI
Part 3

SERVICE CEILING

TWO ENGINE OPERATION AT MAXIMUM CONTINUOUS POWER
ALL CONFIGURATIONS

MODEL: AH-IT(TOW)

DATE: 1 AUGUST 1978
DATA BASIS: FLIGHTTEST

100% ENGINE RPM
65 KIAS

ENGINE: T400-WV-402

FUEL GRADE: JP-4/JP-5

FUEL DENSITY: 6 5/6.8 LB/GAL

Figure 11-10. Service ceiling-chart

11-17

Section XI
Part 4

NAVAIR 01 -H1AAB-1

PART 4 —

BEST RANGE.

The best range charts (figure 11-11, sheets 1 thru 8,

8 TOW missile configuration, twin engine
operation; figure 11-12, sheets 1 thru 8, clean
configuration, twin engine operation; figure 11-19,
sheets 1 thru 4, 8 TOW missile configuration,
single engine operation and figure 11-20, sheets 1
thru 4, clean configuration, single engine
operation) show fuel flow, calibrated airspeed,
torque required and unit range as a function of
gross weight and pressure altitudes for
temperature ranges of -10°C and colder, between -
10°C and +10°C, between +10°C and +25°C, and
+25°C and hotter. Total fuel loading must be
known in order to determine maximum range.

EXAMPLE: Determine the torque required, fuel
flow, airspeed, unit range and the maximum range
of a 12,700 pound helicopter, 8 TOW missile
configuration and twin engine operation at a
pressure altitude of 4000 feet +7.1°C (standard
day). Total fuel loading is 2025 pounds.

Solution:

a. Enter figure 11-11, sheets 3 and 4 at
12,700 pounds gross weight.

b. Move up to 4000 feet pressure altitude
lines, and read:

1. Torque required = 66%Q,

2. Fuel flow = 842 lb/hr,

3. Airspeed = 122 KCAS

4. Unit range = .1517 n mi/lb fuel.

c. To determine maximum range, multiply
the fuel loading by the unit range (2025 x
.1517 = 307 n mi).

RANGE.

The twin engine range charts (figure 11-13, sheets
1 thru 12) and the single engine range charts
(figure 11-21, sheets 1 thru 12) shows performance
data for both 8 TOW missile and clean
configurations. Charts present torque required
and fuel flow (in a tabulated format) and

RANGE

calibrated airspeed and unit range (in graphical
formats) as a function of gross weight and
pressure altitude for temperature ranges of -10°C
and colder, between -10°C and + 10°C, between
+ 10°C and +25°C, and +25°C and hotter. Total fuel
loading must be known in order to determine
range. Since figure 11-13 and figure 11-20 are
identical in format, only one example will be
shown.

EXAMPLE: Determine range and endurance at
cruise condition, MCP, twin engine operation of a
12,700 pound helicopter at a pressure altitude of
4000 feet, +7.1°C (standard day), with a fuel
loading of 2025 pounds for 8 TOW missile
configuration.

Solution:

a. Enter figure 11-13 sheet 4 at 12,700 pound
gross weight. Move up to 4000 feet lines
and read:

8 TOW missile configuration = 131 KCAS.

b. Convert calibrated airspeed to true
airspeed by using the density altitude
chart (figure 11-2 example).

8 TOW missile configuration = 137.2
KTAS.

c. Enter table at top of figure 11-13 sheet 4 at
4000 feet pressure altitude and read 934
lb/hr fuel flow.

d. Divide fuel loading by fuel flow to
determine endurance (2025 -r 934) = 2.17
hours.

e. Multiply KTAS times endurance to
obtain range.

8 TOW missile configuration (137.2 x
2.17) = 297.7 n mi.

TIME AND RANGE VS FUEL.

The time and range vs fuel chart (figure 11-14)
shows the enroute time and the distance that the
helicopter can cover while in level cruise with calm
winds. The only information needed is the cruise
fuel, the fuel flow and the cruise true airspeed.

11-18

NAVAIR 01 -HI AAB-1

Section XI
Part 4

EXAMPLE: Find the time enroute and range
covered while the helicopter consumes 2025
pounds of fuel at a rate of 842 pounds per hour
while cruising at a true airspeed of 122 KTAS.

Solution:

a. Enter figure 11-14 on the upper left at
2025 pounds of fuel. Move horizontally
and interpolate for 842 pounds per hour
fuel flow.

b. Drop to the time scale and read 150
minutes enroute time.

c. Continue to drop and interpolate between
the true airspeed lines to 122 KTAS. Then
project left and read 305 nautical miles
range.

11-19

Section XI
Part 4

NAVAIR 01 -Hi AAB-1

11-20

BEST RANGE

OAT COLDER THAN -10°C

TWO ENGINE OPERATION AT LONG RANGE CRUISE
8 TOW MISSILE CONFIGURATION

MODEL: AH-IT(TOW) 100% ENGINE RPM ENGINE T400-WV-402

DATE: 1 AUGUST 1978 FUEL GRADE JP-4/JP-5

DATA BASIS: FLIGHT TEST FUEL DENSITY: 6 5/6 8 LB/GAL

NAVAIR 01-H1AAB-1

Section XI
Part 4

Figure 11-11. Best range (8 TOW missile configuration) chart (Sheet 2 of 8)

GROSS WEIGHT - LB

.12

8000

9000

10000

11000

12000

13000

14000

11-21

Section XI
Part 4

NAVAIR 01 -HI AAB-1

BEST RANGE

OAT BETWEEN -10°C AND +10°C

TWO ENGINE OPERATION AT LONG RANGE CRUISE
8 TOW MISSILE CONFIGURATION

MODEL: AH-IT(TOW) 100% ENGINE RPM ENGINE: T400-WV-402

DATE: 1 AUGUST 1978 FUEL GRADE: JP-4/JP-5

DATA BASIS: FLIGHT TEST FUEL DENSITY: 6 5/6 8 LB/GAL

8000

9000

10000

11000

GROSS WEIGHT - LB

12000

13000

14000

Figure 11-11. Best range (8 TOW missile configuration) chart (Sheet 3 of 8)

11-22

NAVAIR 01 -HIAAB-1

Section XI
Part 4

11-23

Section XI
Part 4

NAVAIR 01 -HIAAB-1

BEST RANGE

OAT BETWEEN +10° AND +25°C

TWO ENGINE OPERATION AT LONG RANGE CRUISE
STOW MISSILE CONFIGURATION

MODEL: AH-IT(TOW) 100% ENGINE RPM ENGINE T400-WV-402

DATE: 1 AUGUST 1978 FUEL GRADE JP-4/JP-5

DATA BASIS: FLIGHT TEST FUEL DENSITY: 6 5/6 8 LB/GAL

600“

500-

8000

9000

-*-12

10000 11000

GROSS WEIGHT - LB

i .

l t 4 ■

f.....

—

• ■

I 2 -

► •

•

;

•6-

[ j

r ‘

« <

■

12000

13000

14000

Figure 11-11. Best range (8 TOW missile configuration) chart (Sheet 5 of 8)

11-24

NAVAIR 01 -HI AAB-1

Section XI
Part 4

11-25

Section XI
Part 4

NAVAIR 01-H1AAB-1

BEST RANGE

OAT HOTTER THAN +25°C

TWO ENGINE OPERATION AT LONG RANGE CRUISE
8 TOW MISSILE CONFIGURATION

MODEL: AH-IT(TOW) 100% ENGINE RPM ENGINE T400-WV-402

DATE: 1 AUGUST 1978 FUEL GRADE JP-4/JP-5

DATA BASIS: FLIGHT TEST FUEL DENSITY: 6 5/6 8 LB/GAL

(f)

H-

1000 “

900-

800-

700-

600-

500-

w6 ss4^ U | ■ ■ '

8000

9000

10000 11000

GROSS WEIGHT • LB

12000

13000

14000

Figure 11-11. Best range (8 TOW missile configuration) chart (Sheet 7 of 8)

11-26

NAVAIR 01 -HI AAB-1

Section XI
Part 4

11-27

Section XI
Part 4

NAVAIR 01 -HIAAB-1

11-28

NAVAIR 01 -HI AAB-1

Section XI
Part 4

11-29

Section XI NAVAIR 01 -HI AAB-1

Part 4

BEST RANGE

OAT BETWEEN -10°C AND +10°C

TWO ENGINE OPERATION AT LONG RANGE CRUISE
CLEAN CONFIGURATION

MODEL: AH-IT(TOW) 100% ENGINE RPM ENGINE T400-WV-402

DATE: 1 AUGUST 1978 FUEL GRADE JP-4/JP-5

DATA BASIS: FLIGHT TEST FUEL DENSITY 6 5/6 8 LB/GAL

8000

9000

10000

11000

GROSS WEIGHT - LB

12000

13000

14000

Figure 11-12. Best range (clean configuration) chart (Sheet 3 of 8)

11-30

NAVAIR 01 -HIAAB-1

Section XI
Part 4

11-31

Section XI
Part 4

NAVAIR 01 -HI AAB-1

BEST RANGE

OAT BETWEEN +10° AND +25°C

TWO ENGINE OPERATION AT LONG RANGE CRUISE

MODEL: AH-IT (TOW)

DATE: 1 AUGUST 1978
DATA BASIS: FLIGHT TEST

CLEAN CONFIGURATION
100% ENGINE RPM

ENGINE T400-WV-402

FUEL GRADE JP-4/JP-5

FUEL DENSITY 6 5/6 8 LB/GAL

8000

9000

10000 11000

GROSS WEIGHT - LB

12000

13000

14000

Figure 11-12. Best range (clean configuration) chart (Sheet 5 of 8)

11-32

NAVAIR 01 -HIAAB-1

Section XI
Part 4

BEST RANGE

OAT BETWEEN +10° AND +25°C

TWO ENGINE OPERATION AT LONG RANGE CRUISE
CLEAN CONFIGURATION

MODEL: AH-1T (TOW) 100% ENGINE RPM ENGINE T400-WV-402

DATE: 1 AUGUST 1978 FUEL GRADE: JP-4/JP-5

DATA BASIS: FLIGHT TEST FUEL DENSITY 6 5/6 8 LB/GAL

Figure 11-12. Best range (clean configuration) chart (Sheet 6 of 8)

11-33

Section XI
Part 4

NAVAIR 01 -HIAAB-1

BEST RANGE

OAT HOTTER THAN +25°C

TWO ENGINE OPERATION AT LONG RANGE CRUISE

MODEL: AH-IT(TOW)

DATE: 1 AUGUST 1978
DATA BASIS: FLIGHTTEST

CLEAN CONFIGURATION
100% ENGINE RPM

ENGINE T400-WV-402

FUEL GRADE JP-4/JP-5

FUEL DENSITY: 6 5/6 8 LB/GAL

500-

8000

9000

pressure

150

PRESSURE ALTITU DE —.10°° FT
. U -SL-—-—

140

120

100

1000

mT'.TUQE

900

PSSVJBE

800

700

600

12 —

10000 11000

GROSS WEIGHT - LB

12000

13000

14000

Figure 11-12. Best range (clean configuration) chart (Sheet 7 of 8)

11-34

NAVAIR 01 -HI AAB 1

Section XI
Part 4

BEST RANGE

OAT HOTTER THAN +25°C

TWO ENGINE OPERATION AT LONG RANGE CRUISE
CLEAN CONFIGURATION

MODEL: AH-IT(TOW) 100% ENGINE RPM ENGINE T400-WV-402

DATE: 1 AUGUST 1978 FUEL GRADE JP-4/JP-5

DATA BASIS: FLIGHT TEST FUEL DENSITY 6 5/6 8 LB/GAL

GROSS WEIGHT - LB

.12

8000

9000

10000

11000

12000

13000

14000

Figure 11-12. Best range (clean configuration) chart (Sheet 8 of 8)

11-35

Section XI
Part 4

NAVAIR 01-H1AAB-1

RANGE AT MAXIMUM CONTINUOUS POWER

MODEL: AH-IT(TOW)

DATE: 1 AUGUST 1978
DATA BASIS: FLIGHT TEST

OAT COLDER THAN -10°C

TWO ENGINE OPERATION
100% ENGINE RPM

ENGINE T400-WV-402

FUEL GRADE JP-4/JP-5

FUEL DENSITY 6 5/6 8 LB/GAL

H-

I-

ALL CONFIGURATIONS

PRESSURE ALT

TORQUE

FUEL FLOW

0 FT

85 0 %Q

1036 LB/HR

2000

84.9

1032

4000

83 5

1009

6000

76.2

931

8000

69.3

858

10000

62.8

781

12000

56.6

716

200

f—| -4..+—j- t |

pressure altitude — 1000

ICLEAN CONFIGURATION!

too

150

lijow missiue~^onfigurationT

100

P *£SSU

r AL TlTUDF :

Ff T ^°°oh

1 -J—L 4

8000

9000

10000 11000 12000
GROSS WEIGHT — LBS

13000

14000

Figure 11-13. Range at maximum continuous power chart (Sheet 1 of 12)

11-36

NAVAIR 01 -HIAAB-1

Section XI
Part 4

Figure 11-13. Range at maximum continuous power chart (Sheet 2 of 12)

11-37

Section XI NAVAIR 01 -HI AAB-1

Part 4

Figure 11-13. Range at maximum continuous power chart (Sheet 3 of 12)

RANGE AT MAXIMUM CONTINUOUS POWER

OAT COLDER THAN -10°C

MODEL: AH-IT(TOW)

DATE: 1 AUGUST 1978
DATA BASIS: FLIGHT TEST

TWO ENGINE OPERATION
CLEAN CONFIGURATION
100% ENGINE RPM

ENGINE T400-WV-402

FUEL GRADE JP-4/JP-5

FUEL DENSITY: 6 5/6 8 LB/GAL

8000

9000

10000 11000 12000
GROSS WEIGHT - LB

13000

14000

PRESSURE ALT ,jude -

WOO

11-38

NAVAIR 01-H1AAB-1

Section XI
Part 4

RANGE AT MAXIMUM CONTINUOUS POWER

OAT BETWEEN -10°C AND +10°C

TWO ENGINE OPERATION
100% ENGINE RPM

MODEL: AH-IT(TOW)

DATE: 1 AUGUST 1978
DATA BASIS: FLIGHT TEST

ENGINE: T400-WV-402

FUEL GRADE JP-4/JP-5

FUEL DENSITY: 6 5/6 8 LB/GAL

200 "

c n

150"

100 -

50-

150-

c n
\-
o
z

100 -

ALL CONFIGURATIONS

PRESSURE ALT

TORQUE

FUEL FLOW

0 FT

85.0 %Q

1052 LB/HR

2000 _

82.4 _

1019

rTooo_

75.3

934 1

6000

68.5

859

8000

61.9

787

10000

55.6

715

12000

49.9

653

pressur e altitu de - 1000

I I I I 1 ' IT

- HCLEAN CONFIGURATIONI

-4-

=fi I

4 t f -

T8TOW MISSILE CONFIGURATION! "

50 ■

8000

9000

10000 11000 1 2000
GROSS WEIGHT — LBS

13000

14000

Figure 11-13. Range at maximum continuous power chart (Sheet 4 of 12)

11-39

Section XI
Part 4

NAVAIR 01 -HI AAB-1

RANGE AT MAXIMUM CONTINUOUS POWER

OAT BETWEEN -10°C AND +10°C

TWO ENGINE OPERATION
8 TOW MISSILE CONFIGURATION

MODEL: AH-1T (TOW) 100% ENGINE RPM ENGINE T400-WV-402

DATE: 1 AUGUST 1978 FUEL GRADE JP-4/JP-5

DATA BASIS: FLIGHT TEST FUEL DENSITY 6 5/6 8 LB/GAL

.12

8000

9000

10000 11000 12000
GROSS WEIGHT - LB

13000

14000

T.JPOO FT

i . —r-SL

t * ; • t ft:?

PRESSURE altitude

Figure 11-13. Range at maximum continuous power chart (Sheet 5 of 12)

11-40

NAVAIR 01-H1AAB-1

Section XI
Part 4

Figure 11-13. Range at maximum continuous power chart (Sheet 6 of 12)

11-41

Section XI
Part 4

NAVAIR 01-H1AAB-1

RANGE AT MAXIMUM CONTINUOUS POWER

OAT BETWEEN +10°C AND +25°C

TWO ENGINE OPERATION
100% ENGINE RPM

ENGINE T400-WV-402
FUEL GRADE JP-4/JP-5
FUEL DENSITY: 6 5/6 8 LB/GAL

ALL CONFIGURATIONS

PRESSURE ALT

TORQUE

FUEL FLOW

0 FT

74.9 %Q

971 LB/HR

2000

69.7

907

4000

64.1

837

6000

58.5

771

8000

52.8

709

10000

47.4

645

12000

42.1

588

GROSS WEIGHT — LBS

Figure 11-13. Range at maximum continuous power chart (Sheet 7 of 12)

MODEL: AH-IT(TOW)

DATE: 1 AUGUST 1978
DATA BASIS: FLIGHT TEST

11-42

NAVAIR 01 -HI AAB-1

Section XI
Part 4

RANGE AT MAXIMUM CONTINUOUS POWER

OAT BETWEEN +10°C AND +25°C

TWO ENGINE OPERATION
8 TOW MISSILE CONFIGURATION

MODEL AH-IT (TOW) 100% ENGINE RPM ENGINE T400 WV-402

DATE: 1 AUGUST 1978 FUEL GRADE JP-4/JP-5

DATA BASIS: FLIGHT TEST FUEL DENSITY 6 5/6 8 LB/GAL

.12

8000

9000

10000 11000 12000
GROSS WEIGHT - LB

13000

14000

Figure 11-13. Range at maximum continuous power chart (Sheet 8 of 12)

11-43

Section XI
Part 4

NAVAIR 01 -HI AAB-1

RANGE AT MAXIMUM CONTINUOUS POWER

OAT BETWEEN +10°C AND +25°C

TWO ENGINE OPERATION
CLEAN CONFIGURATION

MODEL: AH-IT(TOW) 100% ENGINE RPM ENGINE T400 WV 402

DATE: 1 AUGUST 1978 FUEL GRADE JP-4/JP-5

DATA BASIS: FLIGHT TEST FUEL DENSITY 6 5/6 8 LB/GAL

Figure 11-13. Range at maximum continuous power chart (Sheet 9 of 12)

11-44

NAVAIR 01-H1AAB-1

Section XI
Part 4

PRESSURE ALTITUDE

RANGE AT MAXIMUM CONTINUOUS POWER

MODEL: AH-IT (TOW)

DATE: 1 AUGUST 1978
DATA BASIS: FLIGHT TEST

OAT HOTTER THAN +25°C

TWO ENGINE OPERATION
100% ENGINE RPM

ENGINE T400-WV-402

FUEL GRADE JP-4/JP-5

FUEL DENSITY: 6 5/6 8 LB/GAL

ALL CONFIGURATIONS

PRESSURE ALT

TORQUE

FUEL FLOW

0 FT
2000

4000

6000

8000

10000

12000

66.4 %Q

60.2

55.3

50.2

45.0

39.8

34.9

892 LB/HR

833

769

707

648

588

533

200 '

50'
150 ■

LLI

100 -

f^RESSURE ALTITUDE — 1000 FT
-SLj

■ 1 j y™ '

4CLEAN CONFIGURATION!

8 TOW MISSILE CONFIGURATION]

8000

9000

10000 11000 12000
GROSS WEIGHT — LBS

13000

14000

Figure 11-13. Range at maximum continuous power chart (Sheet 10 of 12)

11-45

Section XI
Part 4

NAVAIR 01 -HIAAB-1

Figure 11-13. Range at maximum continuous power chart (Sheet 11 of 12)

RANGE AT MAXIMUM CONTINUOUS POWER

OAT HOTTER THAN +25°C

TWO ENGINE OPERATION
8 TOW MISSILE CONFIGURATION

MODEL: AH-1T (TOW) 100% ENGINE RPM ENGINE T400-WV-402

DATE: 1 AUGUST 1978 FUEL GRADE JP-4/JP-5

DATA BASIS: FLIGHT TEST FUEL DENSITY 6 5/6 8 LB/GAL

.12

8000

9000

10000 11000 12000

GROSS WEIGHT - LB

13000

14000

PRESSURE ALT| T (jD£ _ 1000 FT

11-46

NAVAIR 01-H1AAB-1

Section XI
Part 4

11-47

Section XI NAVAIR 01 -HI AAB-1

Part 4

Figure 11-14. Time and range vs fuel chart

TIME & RANGE VS FUEL

DATA BASIS: CALCULATED DATA

11-48

NAVAIR 01-H1AAB-1

Section XI
Part 5

PART 5 — ENDURANCE

MAXIMUM ENDURANCE.

The maximum endurance charts (figure 11-15,
sheets 1 thru 4, twin engine operation and figure
11-22, sheets 1 thru 4, single engine operation)
present torque required, calibrated airspeed and
fuel flow as a function of gross weight and pressure
altitude for temperature ranges of -10°C and
colder, between -10°C and +10°C, between +10°C
and +25°C, and +25°C and hotter. Maximum
endurance can be determined if gross weight and
fuel loading are known.

EXAMPLE: Determine the minimum fuel flow,
airspeed, torque required and maximum
endurance for a 12,700 pound helicopter, at 4000
feet +7.1°C (standard day), twin engine operation
and a fuel loading of 2025 pounds.

Solution:

a. Enter figure 11-15 sheet 2 at 12,700
pounds gross weight.

b. Move up to 4000 feet density altitude
lines, and read:

1. Minimum fuel flow required = 626
lb/hr.

2. Minimum airspeed required = 70
KCAS.

3. Minimum torque required = 40%Q.

c. Determine maximum endurance by
dividing total fuel load by minimum fuel
flow required (2025 -r 626 = 3.23 hours).

HOVERING ENDURANCE.

The hover endurance chart (figure 11-16) is shown
for out of ground effect at pressure altitudes of sea
level, 4000 feet, 8000 feet and 12,000 feet for various
gross weights and outside air temperatures. Hover
endurance can be determined if gross weight and
fuel loading are known.

EXAMPLE: Determine the fuel flow and
endurance when hovering OGE at 4000 feet,
+7.1°C (standard day) in a 12,700 pound gross
weight helicopter, twin engine operation with a
fuel loading of 2025 pounds.

Solution:

a. Enter figure 11-6 at 12,700 pound gross
weight. Move right to 4000 feet pressure
altitude line.

b. Drop down to OAT baseline. Move up,
following the trend of the guidelines, to
+7.1°C OAT.

c. Move down, read 996 lb/hr fuel flow.

d. Divide total fuel load by fuel flow to

calculate hover endurance (2025 996 =

2.03 hours).

11-49

Section XI NAVAIR 01-H1AAB-1

Part 5

MAXIMUM ENDURANCE — TWO ENGINE OPERATION

OAT COLDER THAN -10°C

MODEL: AH-IT (TOW)

DATE: 1 AUGUST 1978
DATA BASIS: FLIGHT TEST

ALL CONFIGURATIONS
100% ENGINE RPM

ENGINE T400-WV-402

FUEL GRADE JP-4/JP-5

FUEL DENSITY 6 5/6 8 LB/GAL

700*

600 -•

500 ■

400-

300 ■

8000

9000

10000 11000 12000
GROSS WEIGHT - LB

13000

14000

Figure 11-15. Maximum endurance chart (Sheet 1 of 4)

11-50

NAVAIR 01-H1AAB-1

Section XI
Part 5

11-51

Section XI
Part 5

NAVAIR 01-H1AAB-1

11-52

NAVAIR 01 -HI AAB-1

Section XI
Part 5

11-53

Section XI
Part 5

NAVAiR 01 -HIAAB-1

HOVERING ENDURANCE - TWO ENGINE OPERATION

ALL CONFIGURATIONS
OUT OF GROUND EFFECT

MODEL: AH-IT(TOW) 100% ENGINE RPM ENGINE: T400-WV-402

DATE: 1 AUGUST 1978 FUEL GRADE: JP-4/JP-5

DATA BASIS: FLIGHT TEST FUEL DENSITY: 6,5/6 8 LB/GAL

Figure 11-16. Hover endurance chart

11-54

NAVAIR 01-H1AAB-1

Section XI
Part 6

PART 6 — EMERGENCY OPERATION

SINGLE ENGINE MAXIMUM GROSS
WEIGHT FOR HOVERING.

Refer to maximum gross weight for hovering text
in part 2 (page 10).

SINGLE ENGINE CLIMB

PERFORMANCE.

Refer to climb performance text in part 3 (page 15).

SINGLE ENGINE BEST RANGE.

Refer to best range text in part 4 (page 18).

SINGLE ENGINE RANGE.

Refer to range text in part 4 (page 18).

SINGLE ENGINE MAXIMUM
ENDURANCE.

Refer to maximum endurance text in part 5 (page
49).

SINGLE ENGINE SERVICE CEILING.

Refer to service ceiling text in part 3 (page 15).

ABILITY TO MAINTAIN FLIGHT ON ONE
ENGINE.

This chart (figure 11-24) presents both 8 TOW
missile and clean configurations for lines of gross
weight as a function of pressure altitude and
calibrated airspeed at outside air temperatures of
-20°C, 0°C, +20°C and +40°C.

EXAMPLE: Determine the minimum and the
maximum airspeed (in terms of KCAS) for a 12,000
pound helicopter, 8 TOW missile configuration,
with one engine inoperative at an altitude of 2000
feet, standard day (+11°C).

Solution:

Interpolation between 0°C OAT and +20°C OAT
charts is necessary in order to satisfy the standard
day condition.

a. Enter figure 11-24 at 2000 feet pressure
altitude for both 0°C and +20°C charts.
Read minimum and maximum calibrated
airspeeds for each temperature:

0°C OAT +20°C OAT

Minimum 38 KCAS 50 KCAS

Maximum 98 KCAS 80 KCAS

b. Divide actual temperature by the &
temperature to obtain interpolation
factor (+11°C +20°C = 55%).

c. Interpolate for +11 °C minimum and
maximum calibrated airspeeds using
55% interpolation factor.

1. Minimum airspeed at 11°C, 2000 ft =
44.6 KCAS

2. Maximum airspeed at 11°C, 2000 ft =
88.1 KCAS.

MINIMUM AIRSPEED FOR FLIGHT WITH
ONE ENGINE.

The minimum airspeed for flight with one engine
chart (figure 11-25) presents gross weight as a
function of calibrated airspeed and outside air
temperature for sea level and out of ground effect
condition.

EXAMPLE: Determine the minimum airspeed for
a 12,000 pound helicopter operating on one engine
at sea level, +11°C.

11-55

Section XI
Part 6

NAVAIR 01 -HI AAB-1

Solution:

a. Enter figure 11-25 at 12,000 pound gross
weight. Move left and interpolate
between +10°C and +15°C for +11 °C.

b. Drop down and read 38 KCAS.

11-56

NAVAIR 01-H1AAB-1

Section XI
Part 6

11-57

Section XI
Part 6

NAVAIR 01 -HI AAB-1

11-58

NAVAIR 01-H1AAB-1

Section XI
Part 6

Figure 11-18. Single engine climb performance chart

SINGLE ENGINE CLIMB PERFORMANCE

MODEL: AH-IT(TOW)

DATE: 1 AUGUST 1978
DATA BASIS: FLIGHT TEST

INTERMEDIATE RATED POWER
ALL CONFIGURATIONS
100% ENGINE RPM
65 KIAS

ENGINE T400-WV-402

FUEL GRADE: JP-4/JP-5

FUEL DENSITY: 6.5/6.8 LB/GAL

200

100

11-59

Section XI
Part 6

NAVAIR 01-H1AAB-1

11-60

NAVAIR 01-H1AAB-1

Section XI
Part 6

11-61

Section XI
Part 6

NAVAIR 01-H1AAB-1

11-62

NAVAIR 01-H1AAB-1

Section XI
Part 6

SINGLE ENGINE BEST RANGE

(AT LONG RANGE CRUISE)

OAT HOTTER THAN +25°C

MODEL: AH-IT(TOW) 8 TOW MISSILE CONFIGURATION ENGINE T400-WV-402

DATE: 1 AUGUST 1978 100% ENGINE RPM FUEL GRADE JP-4/JP-5

DATA BASIS: FLIGHT TEST FUEL DENSITY: 6 5/6 8 LB/GAL

Figure 11-19. Single engine best range (8 TOW missile configuration) chart (Sheet 4 of 4)

11-63

Section XI
Part 6

NAVAIR 01-H1AAB-1

11-64

NAVAIR 01 -HIAAB-1

Section XI
Part 6

11-65

Section XI
Part 6

NAVAIR 01-H1AAB-1

11-66

NAVAIR 01-H1AAB-1

Section XI
Part 6

SINGLE ENGINE BEST RANGE

(AT LONG RANGE CRUISE)

OAT HOTTER THAN +25°C

MODEL: AH-IT(TOW) CLEAN CONFIGURATION ENGINE T400 WV 402

DATE: 1 AUGUST 1978 100% ENGINE RPM FUEL GRADE JP-4/JP-5

DATA BASIS: FLIGHT TEST FUEL DENSITY 6 5/6 8 LB/GAL

Figure 11-20. Single engine best range (clean configuration) chart (Sheet 4 of 4)

11-67

Section XI NAVAIR 01-H1AAB-1

Part 6

MODEL: AH-IT(TOW)

DATE 1 AUGUST 1978
DATA BASIS: FLIGHT TEST

SINGLE ENGINE RANGE CHART

OAT COLDER THAN -10°C

INTERMEDIATE RATED POWER
100% ENGINE RPM

ENGINE T400-WV 402

FUEL GRADE JP-4/JP-5

FUEL DENSITY 6.5/6 8 LB/GAL

150-

ULi

I—

100 "

50"

150-

I—

100 -

50"

—

ALL CONFIGURATION

PRESSURE ALT

TORQUE

FUEL FLOW

0 FT
2000

4000

6000

8000

10,000

12,000

51 5%Q

48.4

45.1

41.7

38.3

34.9

31.6

615 LB/HR

585

546

511

476

434

399

8000

PRESSURE ALTITUDE — 1000 FT

-SL--

jCLEAN CONFIGURATION!

♦

t 4

4 4

PRESSURE ALTITUDE - 1000 FT

~~T frSLr

9000

8 TOW MISSILE CONFIGURATION}

10000 11000

GROSS WEIGHT • LB

12000

13000

14000

Figure 11 -21. Single engine range chart (Sheet 1 of 12)

11-68

NAVAIR 01 -HI AAB-1

Section XI
Part 6

SINGLE ENGINE RANGE CHART

OAT COLDER THAN -10°C

INTERMEDIATE RATED POWER
8 TOW MISSILE CONFIGURATION

MODEL: AH-IT(TOW) 100% ENGINE RPM ENGINE T400-WV-402

DATE: 1 AUGUST 1978 FUEL GRADE JP-4/JP-5

DATA BASIS: FLIGHT TEST FUEL DENSITY: 6 5/6 8 LB/GAL

Figure 11 -21. Single engine range chart (Sheet 2 of 1 2)

11-69

Section XI
Part 6

NAVAIR 01 -HI AAB-1

Figure 11-21. Single engine range chart (Sheet 3 of 12)

SINGLE ENGINE RANGE CHART

OAT COLDER THAN -10°C

INTERMEDIATE RATED POWER
CLEAN CONFIGURATION

MODEL: AH-IT(TOW) 100% ENGINE RPM ENGINE T400-WV-402

DATE: 1 AUGUST 1978 FUEL GRADE JP-4/JP-5

DATA BASIS: FLIGHT TEST FUEL DENSITY 6 5/6 8 LB/GAL

8000 9000 10000 11000 12000 13000 14000

GROSS WEIGHT - LB

11-70

NAVAIR 01-H1AAB-1

Section XI
Part 6

SINGLE ENGINE RANGE CHART

OAT BETWEEN -10°C AND +10°C

INTERMEDIATE RATED POWER
100% ENGINE RPM

ENGINE: T400-WV-402
FUEL GRADE: JP-4/JP-5
FUEL DENSITY: 6.5/6 8 LB/GAL

MODEL: AH-IT(TOW)

DATE: 1 AUGUST 1978
DATA BASIS: FLIGHT TEST

1 50 1

CL
CO
OC

LLI

100 -

150-

ALL CONFIGURATION

PRESSURE ALT

TORQUE

FUEL FLOW

0 FT

49.4%Q

600 LB/HR

2000

45.9

563

4000

42.4

521

6000

38.9

483

8000

35.5

446

10,000

32.2

406

12,000

29.0

372

! I I I L J 1 | I"

PRESSURE ALTITUDE - 1000 FT

PRESSURE ALTITUDE — 1000
■S L: =

8000

M ! I i i 1 1

(CLEAN CONFIGURATION!

-.- ■ j

--j

—j

[ . j

—j

j—j

- T

L_]

]

—

}

L—.

8 TOW MISSILE CONFIGURATION}

9000

10000 11000

GROSS WEIGHT - LB

12000

13000

14000

Figure 11 -21. Single engine range chart (Sheet 4 of 12)

11-71

Section XI
Part 6

NAVAIR 01-H1AAB-1

SINGLE ENGINE RANGE CHART

OAT BETWEEN -10°C AND +10°C

INTERMEDIATE RATED POWER
8 TOW MISSILE CONFIGURATION

MODEL: AH-IT(TOW) 100% ENGINE RPM ENGINE T400-WV-402

DATE: 1 AUGUST 1978 FUEL GRADE JP 4/JP-5

DATA BASIS: FLIGHT TEST FUEL DENSITY 6 5/6 8 LB/GAL

Figure 11 -21. Single engine range chart (Sheet 5 of 12)

11-72

NAVAIR 01 -HI AAB-1

Section XI
Part 6

11-73

Section XI
Part 6

NAVAIR 01-H1AAB-1

SINGLE ENGINE RANGE CHART

OAT BETWEEN +10° AND +25°C

INTERMEDIATE RATED POWER
100% ENGINE RPM

ENGINE: T400-WV-402
FUEL GRADE JP-4/JP-5
FUEL DENSITY: 6.5/6 8 LB/GAL

ALL CONFIGURATION

PRESSURE ALT

TORQUE

FUEL FLOW

0 FT

45.1%Q

561 LB/HR

2000

41.7

523

4000

38.4

482

6000

35.1

445

8000

31.8

409

10.000

28.7

372

12.000

25.7

340

Figure 11-21. Single engine range chart (Sheet 7 of 12)

MODEL: AH-IT(TOW)

DATE: 1 AUGUST 1978
DATA BASIS: FLIGHT TEST

11-74

NAVAIR 01 -HIAAB-1

Section XI
Part 6

SINGLE ENGINE RANGE CHART

OAT BETWEEN +10° AND +25°C

INTERMEDIATE RATED POWER
8 TOW MISSILE CONFIGURATION

MODEL: AH-IT(TOW) 100% ENGINE RPM ENGINE T400-WV-402

DATE: 1 AUGUST 1978 FUEL GRADE JP-4/JP-5

DATA BASIS: FLIGHT TEST FUEL DENSITY 6 5/6 8 LB/GAL

GROSS WEIGHT - LB

Figure 11 -21. Single engine range chart (Sheet 8 of 1 2)

11-75

Section XI
Part 6

NAVAIR 01 -HIAAB-1

SINGLE ENGINE RANGE CHART

OAT BETWEEN +10° AND +25°C

INTERMEDIATE RATED POWER
CLEAN CONFIGURATION

MODEL: AH-IT (TOW) 100% ENGINE RPM ENGINE T400 WV-402

DATE: 1 AUGUST 1978 FUEL GRADE JP-4/JP-5

DATA BASIS: FLIGHT TEST FUEL DENSITY 6 5^68 LB/GAL

Figure 11-21. Single engine range chart (Sheet 9 of 12)

11-76

NAVAIR 01 -HIAAB-1

Section XI
Part 6

SINGLE ENGINE RANGE CHART

OAT HOTTER THAN +25°C

INTERMEDIATE RATED POWER
100% ENGINE RPM

ENGINE T400-WV-402
FUEL GRADE: JP-4/JP-5
FUEL DENSITY: 6 5/6 8 LB/GAL

ALL CONFIGURATION

PRESSURE ALT

TORQUE

FUEL FLOW

0 FT
2000

4000

6000

8000

10,000

12,000

39.1 %Q

36 3

33.4

30.5

27.7

24.9

508 LB/HR
475

438

404

372

339

Figure 11 -21. Single engine range chart (Sheet 10 of 12)

MODEL: AH-IT(TOW)

DATE: 1 AUGUST 1978
DATA BASIS: FLIGHT TEST

11-77

Section XI
Part 6

NAVAIR 01 -HIAAB-1

SINGLE ENGINE RANGE CHART

OAT HOTTER THAN +25°C

INTERMEDIATE RATED POWER
8 TOW MISSILE CONFIGURATION

MODEL: AH-IT(TOW) 100% ENGINE RPM ENGINE T400-WV-402

DATE: 1 AUGUST 1978 FUEL GRADE JP-4/JP-5

DATA BASIS: FLIGHT TEST FUEL DENSITY 6.5/6.8 LB/GAL

f igure 11-21. Single engine range chart (Sheet 11 of 12)

11-78

NAVAIR 01 -HIAAB-1

Section XI
Part 6

MODEL AH-1T (TOW)

DATE 1 AUGUST 1978
DATA BASIS: FLIGHT TEST

SINGLE ENGINE RANGE CHART

OAT HOTTER THAN +25°C

INTERMEDIATE POWER
CLEAN CONFIGURATION
100% ENGINE RPM

ENGINE T400-WV-402

FUEL GRADE JP-4/JP-5

FUEL DENSITY 6 5/6 8 LB/GAL

Figure 11 -21. Single engine range chart (Sheet 1 2 of 12)

11-79

Section XI
Part 6

NAVAIR 01 -HIAAB-1

Figure 11 -22. Single engine maximum endurance chart (Sheet 1 of 4)

11-80

NAVAIR 01-H1AAB-1

Section XI
Part 6

11-81

Section XI
Part 6

NAVAIR 01-H1AAB-1

11-82

NAVAIR 01-H1AAB-1

Section XI
Part 6

SINGLE ENGINE MAXIMUM ENDURANCE

OAT HOTTER THAN +25°C

ALL CONFIGURATIONS
100% ENGINE RPM

MODEL: AH-IT(TOW) ENGINE: T400-WV-402

DATE: 1 AUGUST 1978 FUEL GRADE: JP-4/JP-5

DATA BASIS: FLIGHT TEST FUEL DENSITY: 6 5/6 8 LB/GAL

Figure 11 -22. Single engine maximum endurance chart (Sheet 4 of 4)

11-83

Section XI
Part 6

NAVAIR 01 -HIAAB-1

11-84

NAVAIR 01 -HIAAB-1

Section XI
Part 6

ABILITY TO MAINTAIN FLIGHT ON ONE ENGINE

INTERMEDIATE RATED POWER
100% ENGINE RPM

MODEL: AH-IT(TOW)

DATE: 1 AUGUST 1978
DATA BASIS: FLIGHTTEST

ENGINE: T400-WV-402

FUEL GRADE: JP-4/JP-5

FUEL DENSITY: 6.5/6.8 LB/GAL

15-i

§ K>

cc c.
D b
(/)

(/)

15-

z> 5-
c n
co

I0AT = -20^cl

- ■w

L. \

L .

[ \

(

; N

\ j

L . v

rs*

Y 1

X s

JL-

lL\

—

r .

I i

0 5

►0 11

30 15

15-

10-

OAT = 0 6 Cl

\ ]

__ Q

..._4-.

\^\

\ \

tttI

r~ir 1 1
l i i j

' tr-

\ I

. » .

\ '

yr r . 1 . !

1 ,

\ 0

\ °o \

\{ . f . !

\ \

U \

L O

VAN

Vl V

\ \

\ V \

hj\

W\\

>51

i (

100

150

10,

AT =

-- +

>0°C

JjL

W J

p n-

fmmm

" ^

\?N

s, '

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-CLEAN CONFIGURATION

- 8 TOW MISSILE CONFIGURATION

15-|

IQ-

•AT

L±40^(

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„ \

\ i

V N

—

°c

fvj

L 1

\ 1
A

L. ™

50 100

CALIBRATED AIRSPEED — KNOTS

150 0

50 100

CALIBRATED AIRSPEED — KNOTS

150

Figure 11 -24. Ability to maintain flight on one engine chart

11-85

Section XI
Part 6

NAVAIR 01 -HIAAB-1

MINIMUM AIRSPEED FOR FLIGHT WITH ONE ENGINE

SEA LEVEL OUT OF GROUND EFFECT

MODEL: AH-IT(TOW)

DATE: 1 AUGUST 1978
DATA BASIS: FLIGHT TEST

INTERMEDIATE RATED POWER
ALL CONFIGURATIONS
100% ENGINE RPM

ENGINE: T400-WV-402

FUEL GRADE: JP-4/JP-5

FUEL DENSITY: 6.5/6 8 LB/GAL

Figure 11-25. Minimum airspeed for flight with one engine chart

11-86

NAVAIR 01-H1AAB-1

Section XI
Part 7

PART 7 — SPECIAL CHARTS

RADIUS OF TURN AT CONSTANT
AIRSPEED.

The radius of turn at constant airspeed chart
(figure 11-26) presents turn radius as a function of
true airspeed and bank angle.

EXAMPLE: Determine the bank angle and the
turn radius while making a standard 3 degrees per
second turn at an airspeed of 97 KTAS.

Solution:

a. Enter figure 11-26 at 97 KTAS. Move up
and intersect the standard turn line.

b. Read (or interpolate for) bank angle = 15
degrees.

c. Move left and read a turn radius of 3120
feet.

11-87

Section XI
Part 7

NAVAIR 01-H1AAB-1

RADIUS OF TURN AT CONSTANT AIRSPEED

100% ENGINE RPM

MODEL: AH-1T (TOW) ENGINE T400-WV-402

DATE: 1 AUGUST 1978 FUEL GRADE JP-4/JP-5

DATA BASIS: FLIGHTTEST FUEL DENSITY: 6 5/6 8 LB- GAL

Figure 11 -26. Radius of turn at constant airspeed chart

11-88

NAVAIR 01-H1AAB-1 lndex

Acceleration — Climb

INDEX

Page No.
Text Illus

Acceleration G Limitations.

.1-71

1-73

AC Power Supply System.

.1-32

ac armament circuit breaker

panel.

.1-32

1-33

ac power control.

inverter caution lights.

.1-32

inverters switch.

. 1-32

1-33

Advisory Caution and Warning Light

Initial Action.

. .5-2

After Takeoff.

.3-13

Air Capable Ship Operations.

.3-27

launch procedures.

. 3-27

recovery.

. 3-27

stabilized glideslope indicator ....

. 3-27

Airspeed Calibration Chart.

..11-1

11-4

Airspeed Limits.

.. 1-66

Airs tart.

. .5-15

Air Taxi, IFR.

...6-1

Air Taxiing.

. • 3-12

Altimeter Radar AN/APN-17(V)...

• • 7-19

7-21

Armament Circuit Breakers (Pilot) .

.8-11

8-12

Armament Configuration.

. .8-1

Armament Controls and Indicators

(Gunner).

.8-11

8-14

gunner accuracy control panel. . . .

.8-30

8-31

gunner armament control panel. . .

.8-11

8-15

sight hand control (SHC).

. 8-25

8-26

telescopic sight unit (TSU).

. 8-22

8-23

TOW control panel.

. 8-22

8-24

Armament Controls and Indicators

(Pilot).

. . 8-5

8-6

NARCADS .

.. 8-5

8-8

pilot armament circuit breakers . ,

. 8-11

8-12

pilot armament control panel.

.. 8-5

8-7

pilot fixed sight.

.. 8-5

8-10

pilot steering indicator (PSI).

.8-25

8-28

smoke grenade dispenser control

panel.

. 8-11

8-13

Armament Control Panel (Gunner) .

, .8-11

8-15

Armament Control Panel (Pilot)....

.. . 8-5

8-7

Page No.
Text Illus

Armament Firing Modes.

..8-1

Armament Inflight Procedures.

8-37

Armament, Interrelation of .

..8-1

8-2

Armament Switches Cyclic Stick ...

.8-11

Armament System,Wing Stores ....

.8-32

Arrangement Helicopter.

.. 1-2

1-3

1-4

1-5

1-6

Automatic Direction Finder

AN/ARN-83.

.7-24

7-25

Autorotat.ion Characteristics.

. .4-8

Autorotative Landing.

.5-30

Autorotative Landings.

.3-29

autorotation practice.

.3-29

full autorotation landing.

.3-30

hovering autorotation.

.3-30

Auxiliary Fuel System.

.1-27

Battery Overtemp/Thermal

Runaway.

. 5-23

Bearing-Distance-Heading Indicator

(BI)HI).

. 7-20

7-22

Bomb Operation.

8-40

Boost Pump Fuel Failure.

. 5-24

Brake Rotor.

. 1-23

1-21

Canopy Jettison System .

..1-44

1-47

Carrier Qualification.

..3-18

briefing.

carrier qualification and

requalification requirements...

. .3-18

flight scheduling.

..3-18

hanger and flight deck

procedures.

..3-19

Caution Pilot Master System.

.. 1-41

1-42

Ceiling Service.

.11-15

twin engine chart.

11-17

single engine chart.

11-84

Center of Gravity Limitations.

..1-71

1-72

Chip Detector Combining Gearbox

.. 5-27

Chip Detectors Engine.

..1-10

('limb.

..3-13

Index - 1

Index

Climb — Electrical

NAVAIR 01-H1AAB-1

INDEX (Cont)

Page No.
Text Ulus

Climb Performance.11-15

single engine chart.

twin engine chart.

Cold Weather Operation. 6-2

before leaving the helicopter. 6-6

before starting engines. 6-4

engine ground operation. 6-2

engine servicing. 6-2

icing conditions. 6-4

introduction. 6-2

landing. 6.5

main rotor blades and elevators .... 6-4

post flight. 6-6

preparation for flight. 6-3

shutdown. 6-5

starting engines. 6-4

takeoff. 6-4

utilization of manual fuel for cold

starts. 6-3

Combining Gearbox Malfunctions... 5-26
combining gearbox chip

detector.5-27

combining gearbox oil

overtemperature. 5-27

combining gearbox oil pressure

low.5-27

Command Responsibility..3-17

Compass Set AN/ASN- 75 B. 7-10

free gyro operating procedure.7-13

slaved gyro operating procedure... 7-13

Compass Stand-by.1-41

Compressor Stalls.5-20

Control AC Power.1-32

Controls engine. 1-10

Control Panel, Gunner

Accuracy.8-30

Control System Collective.1-34

Control System Cyclic.1-34

Control System Flight.1-34

Control System Malfunctions.5-12

collective control interference.5-12

cyclic control interference.5-12

Control System Tail Rotor.1-34

Copilot/Gunner Seat.1-48

Crew Compartment Doors.1-44

Countermeasures Dispensing System

AN/ALE-39.8-40

Countermeasures System

AN/ALQ-144.8-41

11-59

11-16

7-12

1-8

8-31

1-36

1-6

8-42

Part No.

Text Illus

Cruise.3.13

Cyclic Stick Armament Switches .... 8-11

trigger action. 347

trigger turret fire. 347

wing arm fire. 3.-11

DC Power Supply System.1-28

battery.1-29

dc circuit breaker panel.1-32 1-5

dc power control .1-29

electrical system indicators.1-32

generator.1-29

nonessential bus switch.1-29

Debriefing. 3.28

Defogging/Defrosting.1-48

Density Altitude.ll-l 11.5

Descent. 3.13

Discrepancy Reporting. 3-6

Dispenser Smoke Grenade.8-36 8-34

Ditching. 5.31

ditching power off.5-31

ditching power on.5-31

Doors Crew Compartment. 1-44

Dual Engine Failure During

Takeoff. 5-8

Dummy TSU Ferry Flight.1-11

Dynamic Rollover Characteristics.4-6

Electrical Failure Complete.5-22

Electrical Fire.5-24

Electrical Fire As Evidenced By

Excessive Loads.5-23

Electrical System Malfunctions.5-21

battery overtemp/thermal

runaway.5-23

complete electrical failure.5-23

failure of both generators.5-22

failure of both inverters.5-22

failure of one generator master

arm switch not required.5-21

failure of one generator master
arm switch required.5-21

Index - 2 Change 1

Section I
Part 2

NAVAIR Q1-H1AAB-1

INDEX (Cont)

Index

Electrical — Flight

Page No.
Text Illus

main

Elimination of Smoke and Fumes

emergency egress
rescue .

Emergency Operations
ability to maintain flight one

engine.

minimum airspeed for flight

single engine maximum

scope

Endurance

maximum.

Engine Failure (Dual) During

Takeoff.

Engine Failure (Single) During

airstart

compressor

mt)

.. 5-22

.. 1-37

1-3

,.. 5-23

.... 5-5

,... 5-8

.... 5-7

.... 5-5

5-6

.... 5-5

... 1-41

1-42

... 1-44

...1-43

1-45

... 1-44

...1-41

1-42

... 1-44

..11-55

11-85

..11-55

11-86

..11-55

11-18

..11-55

11-15

..11-55

11-49

..11-55

11-18

..11-55

11-15

....5-1

..11-49

11-54

..11-49

11-50

thru

11-53

.1-1

...5-15

.5-8

.5-7

....5-14

...5-15

.... 5-20

Page No.
Text Illus

Engine Malfunctions (Cont)

dual engine failure.5-16

engine chip detector caution

light.£>20

engine fire in flight.5-19

engine oil overtemperature.5-20

engine oil pressure low.5-20

engine over speed rotor rpm (nr) ... 5-18

engine shutdown in flight.5-15

engine underspeed gas prod (ng) . .5-18

fire both engine in flight.5-19

power turbine governor (nf)

failure.5-18

single engine failure (hige).5-15

single engine failure (hoge).5-15

single engine failure (inflight).5-14

Engine Torque Limits ...1-66

Engine Wash Procedures.1-56

engine desalinization rinse.1-62

engine performance recovery

wash.1-62

Equipment, Personnel Flying.2-3

Exterior Inspection .3-6

Exterior Lights.1-52

anti-collision light .1-52

fuselage formation lights.1-52

rotor tip formation lights.1-52

searchlight .1-52

External Power Receptacle.1-32

Extinguisher Fire.1-44

Ferry Flights, Dummy TSU.1-11

Field Carrier Landing Practice.3-17

night FCLP.3-18

Fire Electrical .5-24

Fire in Flight Fuselage.5-24

Fixed Sight Pilot.8-5

Flare Operation.8-40

Flight Characteristics.

autorotation characteristics.4-7

control feed back.4-1

dynamic rollover characteristics.... 4-5

hovering capability.4-5

maneuvering flight .4-2

1-67

3-7

1-53

1-54

1-55

8-10

4-8

4-3

4-4

Index - 3

Index

Flight — Gunner

NAVAIR 01-H1AAB-1

INDEX (Cont)

Flight Characteristics (Cont)

Page No.
Text Ulus

launch and recovery operations

Flight With Crosstube Fairings

Fuel Filter Engine Impending

fire extinguishers and

Part No.
Text Ulus

Fuel System Malfunctions .

.. 5-25

engine driven fuel pump failure .

..5-25

engine fuel filter impending

bypass.

fuel boost pump failure .

Fuel Supply System.

1-23

... 4-6

crossfeed valve switch .

.. 1-24

1-25

.. 1-34

engine driven fuel pumps.

.. 1-34

1-36

forward and aft fuel boost

1-36

caution lights.

.. 1-37

forward and aft fuel low

.. 1-34

1-36

caution lights .

fuel filter caution lights .

.. 1-27

fuel interconnect valve switch

1-24

1-95

fuel pressure indicator .

.. 1-24

A. CjkJ

1-26

fuel quantity indicator .

1-24

1-26

fuel system caution lights .

.. 1-27

..3-20

fuel switch engine 1 and

.. 3-20

engine 2 .

1-25

..3-20

Functional Checkflight Procedures

. .3-35

checkflights and forms .

. .3-35

1 71

check pilots .

n *1

introduction .

7-3

Fuselage Fire In Flight .

3-31

3-32

. 3-33

Gage Fluid Level Sight .

Gage Hydraulic .

1-34

Gearbox Combining .

1-10

Gearbox Malfunctions 42 Degree

11-9

and 90 Degree .

5-27

1-57

42°/90° chip detector .

1-58

42°/90°gearbox oil over¬

1-59

temperature or low pressure .

.5-27

Generator-Starter.

Gliding Distance From Land ..

2-3

Ground Emergencies.

Ground Operations.

fire guard.

helicopter acceptance.

3-6

preflight inspection.

3-6

1-16

Gun Pod Wing ...

8-33

1-64

Gun Pod Operation, Wing ...

8-39

Gunner Armament Controls and

..5-25

Indicators ....

8-11

8-14

Gunner/Copilot Seat.

. 1-48

1-6

1-13

Guns, Telescopic Sight Unit . .

.8-30

Index-4 Change 1

NAVAIR 01 -HI AAB-1

Index

Harness — Instruments

INDEX (Cont)

Page No.
Text Illus

Harness Shoulder.1-48

Heater Fuel Control Line.1-15

Helicopter .1-1

Helicopter Arrangement.1-2 1-3

thru

1-6

Helicopter Operations On Air

Capable Ships.3-22

launch procedures.3-22

recovery procedures.3-22

Helmet Sight Subsystem (HSS). 8-17 8-18

High Altitude Effects.6-7

Hot Start.5-7

Hot Weather Operations .6-6

desert operations.6-6

preparation for flight .6-6

Hovering Capability.4-6

Hydraulic Malfunctions.5-8

complete (dual) loss of flight

control hydraulic boost.5-9

hydraulic actuator/servo

malfunctions.5-9

hydraulic system no 1 failure.5-8

hydraulic system no 2 failure.5-9

wave-off with complete hydraulic

failure.5-11

Hydraulic Power Supply System .... 1-34

fluid level sight gage.1-34

hydraulic filter and indicator.1-34

hydraulic gage.1-34

hydraulic system 1 and hydraulic

system 2 caution lights.1-34

hydraulic system switch .1-34 1-35

Ice and Rain Removal System.1-48

Icing Conditions.6-4

Indicated Torque Required To
Hover.11-10

Indicator Airspeed.1 : 39

Indicator Combining Gearbox Oil
Temperature and Pressure.1-22

1-51

11-13

11-14

indicator Copilot/Gunner Attitude .. 1-41
Indicator Free Air Temperature.1-41

Indicator Fuel Pressure.1-24 1-26

Indicator Fuel Quantity.1-24 1-26

Indicator Hydraulic Filter.1-34

Indicator Pilot Attitude.1-39

Indicator Vertical Velocity.1-39

Indicators Electrical System .1-32

Indoctrination.2-1

currency.2-3

flight crew designation qualification

and requirements.2-2

ground training.2-1

night and instrument flights.2-3

personnel flying equipment.2-3

pilot flight training.2-1

pilot ground training.2-1

when beyond gliding distance

from land.2-3

Inflight Emergencies.5-8

Instrument Markings.1-66 1-67

Instrument Procedures.6-1

air taxi.6-1

descent.6-2

instrument climb.6-2

instrument cruise flight.6-2

instrument flight checklist.6-1

simulated instrument flight.6-1

start.6-1

Inflight Procedures — All

Armament.8-37

bomb operation.8-40

flare operation.8-40

rocket operation.8-39

smoke grenade dispenser

operation.8-39

TOW operation .8-38

turret operation.8-37

wing gun pod operation.8-39

Instruments .1-37

airspeed indicator.1-39

altimeter.1-39

copilot/gunner attitude indicator .. 1-41

free air temperature indicator.1-41

pilot attitude indicator.1-39

stand-by compass.1-41

vertical velocity indicator.1-39

Index - 5

Index

Instruments — Malfunctions

NAVAIR 01-H1AAB-1

INDEX (Cont)

Page No.

Text

Illus

Instruments and Indicators Engine

. 1-15

1-6

1-8

Intercommunication System

AN/AIC-18.

.7-13

7-14

Interior Inspection Copilot/Gunner.

.. 3-7

Interior Inspection Pilot.

.. 3-8

Interior Lights.

. 1-52

crew compartment lights.

. 1-52

1-54

1-55

pilot, copilot/gunner console

lights.

. 1-56

1-54

1-55

pilot, copilot/gunner instrument

lights.

. 1-52

1-54

1-55

Interrelation of Armament.

..8-1

8-2

Inverter Failure Main.

.5-22

Jettison Canopy System .

. 1-44

147

Jettison Wing Stores.

.5-28

8-32

Landing.

.5-30

autorotative landing.

.5-30

crosswind landing.

3-14

emergencies, landing.

5-30

high speed approach and

landing..

.3-15

maximum gross weight landing . .

.3-15

normal approach and landing .. . .

.3-14

single engine landing.

. 5-30

sliding landing .

.3-15

slope landing.

.3-14

steep approach and landing.

.3-14

Landing Autorotative.

.3-22

Landing Emergencies.

.5-30

Part No.
Text Illus

Lights Forward and Aft Fuel

Boost Caution .

Lights Forward and Aft Fuel

Lights Pilot, Copilot/Gunner

Lights Pilot, Copilot/Gunner
Instrument .

Limitations Rotor Brake

limitations for towing the

Malfunctions Electrical Systems

... 1-37

...5-31

... 1-71

... 1-52

1-53

... 1-34

... 1-52

1-54

1-55

... 1-52

... 1-27

... 1-52

1-53

... 1-52

... 1-32

... 1-56

1-54

1-55

... 1-52

1-54

1-55

... 1-52

1-53

... 1-71

... 1-66

... 1-64

... 1-65

...5-28

... 5-28

5-29

..5-21

.... 4-1

...5-12

... 5-26

... 5-21

...5-14

Index - 6

NAVAIR 01-H1AAB-1

Index

Malfunctions — Power

INDEX (Cont)

Page No.
Text Dlus

Malfunctions Fuel System..5-24

Malfunctions Hydraulic.5-8

Malfunctions Transmission.5-25

Malfunctions 42°/90° Gearbox.5-27

Mast Bumping.4-3,5-13

Markings Instrument.1-66

Maneuvering Flight.4-2

Maximum Gross Weight for
Hovering.11-10

Minimum Crew Requirements.1-66

Mission Planning.3-1

factors affecting helicopter lift

capability.3-1

general precautions ..3-3

introduction .3-1

requirements for mission

planning.3-3

weight limitations applicable to

helicopters.3-2

Mountain and Rough Terrain

Flying ...! .6-6

adverse weather operation.6-9

effects of high altitude.6-7

landing site evaluation.6-7

summary.6-9

turbulent air flying techniques.6-7

wind direction and velocity.6-6

NARCADS. 8-5

NATOPS Evaluation.10-1

concept. 10-1

flight evaluation.10-3

flight evaluation grading criteria.. 10-4

grading instructions.10-2

ground evaluation.10-2

implementation.10-1

open book exam .10-7

report .10-6

Night and Instrument Flights.2-3

Night Flying.3-17

restrictions on night flying.3-17

Night Operations.3-22

helicopter lighting.3-28

postflight procedures.3-28

preflight procedures.3-28

taxi and operations.3-28

1-67

11-11

11-12

3-4

6-11

6-8

6-9

6-11

8-8

Oil Overtemperature Combining

Gearbox .

Oil Pressure Low Combining

Panel AC/Armament Circuit

Particle Separator Engine Air

range

Pilot Armament Controls and

Post Firing/Before Landing

Check—All Armament.

Postflight External Inspection

engine air particle

engine idle stop release switch

Page No.
Text Illus

..5-26

..5-27

..5-20

..1-11

1-14

..3-19

... 1-32

1-33

... 1-41

1-42

... 1-32

1-5

.... 1-7

1-12

...11-1

..11-15

..11-55

..11-49

..11-18

..11-87

...11-1

11-2

..11-10

... 8-5

8-6

... 1-48

1-5

.. 8-25

8-28

.... 4-2

. .8-40

. 3-16D

...3-11

.1-7

1-11

.1-7

1-10

_1-12

,... 1-15

1-8

.1-7

1-12

.... 1-10

1-8

.1-7

1-13

.... 1-12

1-8

.... 1-12

1-8

....1-11

1-14

Change 1 Index - 7

Index

Power — Shaft

NAVAIR 01-H1AAB-1

INDEX (Cont)

Part No.

Text

Illus

Power Settling.

Power Supply AC System.

... 1-32

Power Supply DC System.

... 1-28

Power Supply Hydraulic System .

... 1-34

Pre-Entry Inspection.

.... 3-6

Preflight Procedures (Armament).

. . 8-32

before exterior check — all

armament — preflight.

.. 8-32

exterior check — preflight.

.. .8-32

Pre-Landing Check .

...3-13

Pressure Altitude.

...11-1

Pressure Fueling.

.. . 1-27

1-16

1-64

Pressure Hot Fueling.

... 1-64

emergency shutdown.

... 1-64

fueling personnel .

... 1-63

ground crew requirements.

... 1-64

Pre-Start Checklist.

.... 3-9

Pre-Takeoff Checklist..

...3-11

Procedures (FCF).

... 3-29

before preflight...

...3-29

exterior check.

... 3-30

flight checks..

,.. 3-40

functional check flight.

.. 3-29

hover checks.

.. 3-40

interior inspection (pilot).

..3-35

pre-entry inspection..

..3-35

safety check.

..3-30

shutdown...

start.

Prohibited Maneuvers.

.. 1-66

Pumps Engine Driven Fuel.

.. 1-24

Pylon Rock.

Quick Stop..

Radar Altimeter AN/APN-171(V) ..

.7-19

7-21

Radar Beacon AN/APN-154(V)..

.7-10

7-11

Radar Warning System AN/APR-39 .

. 8-40

8-44

operating procedures. . ..

. 8-41

Page No.
Text Illus

Radio AN/ARC 114A FM.7-1

Radio AN/ARC-159(V)1 UHF.7-5

Range Ch art. 11-18

TOW.

clean

twin engine

single engine

Receptacle External Power.1-32

Rendezvous.8-33

carrier type rendezvous.8-33

running rendezvous.8-33

Requirements.3-29

conditions requiring functional

checkflights.3-29

Rocket Launcher, LAU Series.

Rocket Operation.8-39

Rockets.8.32

Rollover Dynamic Characteristics .... 4-5

Rotor Brake Limitations.1-66

Rotor Brake Pressurized In Flight.. 5-28

Rotor Brake.i_23

operation of rotor brake.1-23

Rotor Droop. 4.7

Rotor Main.i-i6

Rotor System.i-i6

main rotor.

RPM caution system.1-18

tail rotor.i.jg

Rotor Tail .148

SCAS Failure.5.11

Scheduling..

Seat Pilot.i_4g

Servicing and Fueling.1-56

Shaft Horsepower Versus Torque_11-2

7-3

7-6

11-20

thru

11-27

11-28

thru

11-35

11-36

thru

11-47

11-68

thru

11-79

18-27

1-21

1-17

1-21

1-5

1-57

1-58

I- 59

II - 6
11-7
11-8

Index-8 Change 1

NAVAIR 01-H1AAB-1

Index

Ships — Transmission

INDEX (Cont)

Page No.
Text Illus

Ship Based Procedures.

.3-17

Shore Based Procedures.

.3-5

introduction .

.3-5

Shoulder Harness.

.1-48

Shutdown.

... . 3-16D

Sight Hand Control (SHC)-

. 8-25

8-26

Sight Subsystem Helmet (HSS)

.8-17

8-18

Single Engine Failure During
Takeoff.

.5-7

Single Engine Landing.

.5-30

Skid Tail .

.1-37

Smoke and Fumes in Cockpit
Elimination of.

.5-23

Smoke Grenade Dispenser.

.8-32

8-34

control panel .

_8-11

8-13

operation .

8-11,8-39

Special Procedures.

. 3-29

Speed Range.

. 1-2

Stability and Control Augmentation

System (SCAS) .

.1-37

control panel.

.1-37

description.

.1-37

SCAS (SAS) release switch .

.1-37

1-38

Stabilization Electronic Control
Amplifier.

_8-22

Start Checklist.

.3-9

Start Switch.

.1-15

engine instruments and
indicators.

.1-15

1-6

fuel control line heater.

.1-15

1-8

starter-generator.

.1-15

Starter Limitations .

.1-66

Switch Crossfeed Valve.

.1-24

1-25

Switch Engine RPM.

.1-12

1-8

Switch Fuel Engine 1 and
Engine 2.

.1-24

1-25

Switch Fuel Interconnect Valve

.1-24

l-2£

Switch Hydraulic System.

.1-34

1-35

Switch Inverters.

.1-32

1-33

Switch Nonessential Bus.

.1-29

Switch release engine idle stop

.1-12

1-8

Switch SCAS (SAS) release ...

.1-37

1-38

Switch Start.

.1-15

Synchronized Elevator.

.1-37

1-3

TACAN AN/ARN-84(V) .

.7-20

7-23

Page No.

anti-torque malfunction wnne
pt p hover .

complete loss of tail rotor

stabilization electronic control

Transmission Malfunctions,

Text

Illus

. 7-20

7-20B

7-20A

. .4-2

. . . 9-2

. .9-2

..5-12

..5-13

,..5-12

...5-13

... 1-37

.... 5-7

.... 1-2

. . . 6-4

. . 3-12

.. 8-22

8-23

.. 8-30

. . 8-30

. . 8-32

..11-18

11-48

... 1-66

1-67

8-24

. . 8-19

. . 8-25

. . 8-19

. . 8-25

8-28

. . 8-25

8-26

. . 8-22

. . 8-29

. . 8-22

8-23

. . 8-22

8-24

. . 8-20

. . 8-25

8-21

... 8-38

... 1-18

1-19

. . 5-25

Change 1 Index - 9

Index 10

Transmission — Wing

NAVAIR 01-H1AAB-1

INDEX (Cont)

Page No.

Part No.

Text

Illus

Text

Illus

Transmission System.

..1-18

combining gearbox oil

temperature and pressure

Ventilating System.

indicator.

.. 1-22

.1-48

1-49

main rotor transmission system .

.. 1-18

1-19

defrosting/defogging.

.1-48

tail rotor transmission system ...

.. 1-22

environmental control unit

Transmission Tail Rotor System ..

.. 1-22

(ECU).

.1-52

1-51

Transponder Identification

rain and ice removal system.

.1-48

1-51

AN/APX-72.

. 7-13

7-16

ventilating system operation.

.1-48

thru

Vibration Identification.

. .4-7

Turbulent Air Flying Techniques ..

7-18

Voice Security System TSEC/KY-28

..7-1

7-3

.. 6-7

6-8

6-10

6-11

Truck and Crew.

Turn at Constant Airspeed.

.11-87

11-88

Turret Operation.

. 8-37

Turret System GTK4A/A.

.. 8-3

Wash Procedures, Engine.

.1-56

functions.

8-4

Waveoff.

.3-16

Types of Takeoff.

autorotative approach.

.3-16

confined area takeoff.

..3-13

power-on approach.

.3-16

crosswind takeoff.

..3-13

Weapons Replaceable Assemblies . . .

8-20

maximum power takeoff.

..3-13

Weapon System.

..8-1

normal takeoff from hover.

.3-12

introduction .

..8-1

normal takeoff from ground.

,.3-12

Weight Takeoff Gross.

..1-2

normal takeoff to hover.

..3-12

Wind Direction and Velocity.

..6-6

sliding takeoff.

..3-12

Wind Envelope .

3-22

takeoff performance..

..3-12

Wind Limitations .

3-21

Wing Gun Pod.

.8-32

8-33

Wing Gun Pod Operation.

8-39

Wing Stores Armament System ....

.8-32

Wing Stores Jettison.5-28, 8-32

copilot/gunner procedures for

jettisoning.

.5-28

pilot procedures for jettisoning ...

.5-28

UHF Direction Finder AN/ARA-50.

.. 7-5

7-9

UHF Radio AN/ARC 159(V)1.

Uncommanded Right Roll During

..7-5

7-6

Flight Below 1G.

. 5-14

Index -10 Change 1

NAVAIR 01 -HI AAB-1

FO-O Table of
Contents

Table of Contents

Fuel Schematic Diagram.FO-1

Hydraulic Schematic Diagram.FO-2

Flight Control System.FO-3

Electrical Schematic Diagram.FO-4

SCAS and Flights Controls.FO-5

Pilot Cockpit Diagram.FO-6

Copilot/Gunner Cockpit Diagram.FO-7

Interrelation of Armament..FO-8

Weapons Replaceable Assemblies.FO-9

☆U.S. GOVERNMENT PRINTING OFFICE; 1983 - 639 - 009/1016

FO-O

Change 1 Reverse blank

□ VENT

FUEL PRESSURE

| -» | CHECK VALVE

IHI

SHUTOFF VALVE

DRAIN VALVE

MAIN FUEL

NAVAIR 01-H1AAB-1

FO-1. Fuel Schematic Diagram

□

FUEL SUPPLY

AIR PRESSURE

FUEL PRESSURE

QUICK DISCONNECT
AUXILIARY FUEL SYSTEM

210062-50A

FO-1
Reverse blank

SOLENOID VALVE
(2 WAY - 2 POSITION)-

RELIEF VALVE
FULL FLOW 3850 PSID
RESEAT 3250 PSID MIN.-

FILTER MODULE'

SYSTEM LEADING PARTICULARS

HYDRAULIC FLUID: MIL-H-83232
SYSTEM TEMPERATURE RANGE: -65° TO 275°F
SYSTEM OPERATING PRESSURE: 3000 PSIG
SYSTEM NO. 1 CAPACITY: 259 CUBIC INCHES
SYSTEM NO. 2 CAPACITY: 330 CUBIC INCHES
MAXIMUM TEMPERATURE RISE: APPROXIMATELY 75.0

FORE AND AFT (SCAS) ACTUATOR

LATERAL (SCAS) ACTUATOR

GUNNERS CAUTION PANEL

NO 1 NO 2

OC GEN DC GEN

SPARE

ENG 1
CHIP OETR

ENG 2
CHIP OETR

XMSN
Oil PRESS

XMSN C BOX

OIL HOT OIL PRESS

ENG 1
GOV MAN

ENG 2
GOV MAN

AC MAIN

AC STBY

FWD AFT

FUEl. LOW FUEL LOW

ENG 1
FUEL FITfl

ENG 2
FUEl FLTR

NO 1

HYO PRESS

NO 2

HYO PRESS

NO 1

HYD TEMP

NO 2
HYO TEMP

ENG t
OIL PRESS

ENG 2
OIL PRESS

CHIP OETR

CHIP DETR

C BOX
CHIP OETR

XMSN
CHIP OETR

PILOTS CAUTION PANEL

NAVAIR 01-H1AAB-1

FO* 2. Hydraulic Schematic
Diagram

TO SECU

ARMAMENT ARMAMENT

SYSTEM ON SYSTEM OFF

SOLENOID VALVE
SCHEMATIC DIAGRAM

DETAIL A

210076-78

IDLE GOV FORCE SAS

0000

STOP CONT TRIM PWR

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jr~~ y m //'

/' ,"*fy ♦ \\t»\ V ,

X , W/"*' / #X y X

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v./ < ,/

V x '.v / /

,/A - s -T“-SEE DETAIL

ARMAMENT
COMPENSATOR
UNIT-

SENSOR AMPLIFIER UNIT

SCAS CONTROL PANEL

NAVAIR 01-H1AAB-1

FO-3. Flight Control
System

W////A COLLECTIVE
Mill TAIL ROTOR
1 I CYCLIC
ES3 SCAS

N2/83

210060-3

Change 1

FO-3
Reverse blank

LEGEND

I AC LOAD DISTRIBUTION

Sdc LOAD DISTRIBUTION

%^%^jP0WER FROM GENERATOR

POWER FROM BATTERY

POWER FROM -EXTERNAL POWER
relay

TURRET DRIVE MOTORS POWER
TURRET GUN MOTOR POWER

NAVAIR 01-H1AAB-1

FO-4. Electrical Schematic
Diagram

DF ANTENNA ARA-50
TACAN RT ARN-84(V)

MAIN

INVERTER

VOLTMETER

INVERTER

RELAY

115/28V~
AUTO v
TRANSFORMER

STBY

INVERTER

RLY

REF XFMR

BDHI

AC FAILURE RELAY
AC TRANSFORMER (26 VOLTS)

ADF INDICATOR ARN-83
ENGINE VIBRATION METER RECEPTICAL
FUEL QUANTITY INDICATOR
GYROSYN COMPASS, ASN-75B
IFF COMPUTER
RADAR ALTIMETER, APN-171
SCAS POWER
TURRET SIGHT

ADF INDICATOR ARN-83
COMPASS INDICATOR, ASN-75B
ENGINE OIL PRESSURE INDICATOR
FUEL PRESSURE INDICATOR
GEARBOX OIL PRESSURE INDICATOR
TORQUE PRESSURE INDICATOR
TRANSMISSION OIL PRESSURE INDICATOR

DF AMPLIFIER ARA-50
ALTIMETER ENCODER
FORMATION LIGHTS POWER
ROTOR TIP LIGHTS
TACAN ARN-84M

DF AMPLIFIER ARA-50
ECU POWER

MAJN INVERTER CONTROLj
TACAN CONVERTER ARN-84(V)

ATTITUDE SYSTEM
TURRET POWER
SECU POWER
HSS POWER

TOW

POWER

TOW SIGNALS
HSS SIGNALS

ADF RECEIVER, ARN-83
ALTIMETER VIBRATOR
ANTICOLLISION LIGHT
AUXILIARY FUEL SYSTEM
CAUTION LIGHTS
COCKPIT LIGHTS (MAP)

COUNTERMEASURES DISPENSING SYSTEM, ALE-39
COUNTERMEASURES SYSTEM, ALQ-144
DC DUAL VOLTMETER

ENGINE AIR BYPASS VALVE (PARTICLE SEPARATOR)

ENGINE NO. 1 AND NO 2 OIL PRESSURE
ENGINE NO. 1 AND NO 2 ITT
ENGINE START NO 1 AND NO. 2

ENGINE XMSN, AND GEARBOX OIL TEMPERATURE INDICATOR
FIRE DETECTION ENGINE NO 1 AND NO 2
FIRE EXTINGUISHER RESERVE AND MAIN
FM RADIO ARC 114A
FM KY 28

FUEL BOOST FORWARD AND AFT
FUEL HEATER CONTROL

FUEL INTERCONNECT AND CROSSFEED VALVES
FUEL VALVE
FORCE! TRIM

GENERATORS NO 1 AND NO 2 RESET
GOVERNOR CONTROL
GUN POD INBD
GUN POD OUTBD

GUNNER WING STORES JETTISON
HYDRAULIC CONTROL (AND TEST)

ICS GUNNER
ICS PILOT

IDLE STOP SOLENOID

IFF TRANSPONDER APX-72

IFF TRANSPONDER TEST APX-72

INVERTER CONTROL RELAY

MANUAL GOVERNOR

INVERTER STANDBY POWER

MASTER CAUTION LIGHT

NAVIGATION LIGHTS

OVERSPEED GOVERNOR

PILOT AND GUNNER CONSOLE LIGHTS

PILOT AND GUNNER INSTRUMENi LIGHTS

PILOT WING STORES JETTISON

PITOT HEATER

RADAR ALTIMETER, APN-171
RADAR BEACON, APN-154
RADAR WARNING SYSTEM, APR-39
RATE GYRO (TURN AND SLIP)

ROCKET POWER
ROTOR BRAKE LIGHT
RPM CAUTION
SCAS POWER
SEARCHLIGHT CONTROL
SEARCHLIGHT POWER
SMOKE GRENADES
TMS POWER

TRANSMISSION CHIP LIGHT PANEL
TRIPLE TACHOMETER
TURRET ELEVATION STOW
TURRET BUS CONTROL
UHF KY 28
VENT BLOWER
UHF RADIO
WEAPONS CONTROL
WEAPON FIRE
WING STORES JETTISON
WING STORES POWER

N2/83 1
210475-12

FO-4

Change 1 Reverse blank

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NAVAIR 01 -HI AAB-1

FO-6.

Pilot Cockpit
Diagram

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N2/83

210900-142

Change 1

FO-6
Reverse blank

NAVAIR 01-H1AAB-1

FO-7. Copilot/Gunner
Cockpit Diagram

FO-7

Change 1 Reverse blank

N2/83

210900-144

(

NAVAIR 01-H1AAB-1

NAVAIR 01 -HIAAB-1

FO-8. Interrelation
of Armament

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210071-35

FO-8
Reverse blank

NAVAIR 01-H1AAB-1

FO-9 Weapons Replaceable
Assemblies

OPERATOR

TRACKING

VISUAL POSITION
DATA (TSU LOS)

SIGHT HAND
CONTROL

STEERING COMMAND

TOW

CONTROL

PANEL

STABILIZATION CONTROL

AMPLIFIER

• STABILIZATION CIRCUIT

• ERROR SIGNAL RESOLU1

rioN

•MOTION COMPENSATION

•OPEN LOOP STEERING

SERVO

ELECTRONIC

CONTROL

UNIT

AIRCRAFT POWER INPUT
(TURNED ON BY TCP)

ARTICULATED

PYLON

(BOMBRACK)

MISSILE

LAUNCHERS

ELECTRONIC

POWER

SUPPLY

MISSILE
SELECTION

PILOT

STEERING

INDICATOR

PILOT STEERING COMMANDS/
STATUS INDICATORS

MISSILE CONTROL
AMPLIFIER
GUIDANCE COMMANDS
TIMING PROGRAMMER
• BIT

I • m

PRE-FIRE/

FIRE

WIRECUT

FUNCTIONAL ELEMENTS

1 STABILIZES SIGHT

2 CONTROLS AND DISPLAYS

3 INFRARED

4 MISSILE COMMAND

5 LAUNCHER

GUIDANCE COMMANDS/
SELF BALANCE

MISSILE

STATUS

FO-9
Reverse blank

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